T.O. GR1F16CJ1FLIGHT MANUAL
HAF SERIES AIRCRAFTF16C/DBLOCKS 50 AND 52+

LOCKHEED MARTIN CORPORATION
F3365790C2002
F4262001D0058

GR1F-16C-1-0001X

Commanders are responsible for bringing this publication to the attention of all Air Force personnel cleared for
operation of subject aircraft.

15 OCTOBER 2002
CHANGE 115 JUNE 2003

T.O. GR1F16CJ1DISCLOSURE STATEMENT
This information is furnished on the conditions that it will not be released to another nation without the
specific authorization of the Department of the Air Force of the United States; it will be used for military
purposes only; the recipient will report promptly to the United States any known or suspected compromise;and the information will be provided substantially the same degree of security afforded it by the Departmentof Defense of the United States. Also, regardless of any other markings on the document, it will not be
downgraded or declassified without the written approval of the originating US agency. Any request for thisdocument should be referred to OOALC/YPXG, 6089 Wardleigh Road, Hill AFB, UT 840565838.
SUPPLEMENTAL NOTICEThis manual is incomplete without T.O. GR1F16CJ11 and T.O. GR1F16CJ12.
SUPERSEDURE NOTICESee Technical Order Index, T.O. GR0171, for current status of Flight Manuals, Safety Supplements, Opera
tional Supplements, and Flight Crew Checklists.
This manual supersedes T.O. 11C130 and Interim Safety Supplements T.O. GR1F16CJ1SS27 and T.O.
GR1F16CJ1SS29, and Interim Operational Supplement T.O. GR1F16CJ1S31.

T2Change 1

T.O. GR1F16CJ1INSERT LATEST CHANGED PAGES. DESTROY SUPERSEDED PAGES.

LIST OF EFFECTIVE PAGESNOTE:Changes to the current issue are indicated by a vertical line in the outer margins of the page.
Dates of issue for original and changed pages are:Original . . . 0 . . . . 15 Oct 02

Change . . . . 1 . . . . 15 Jun 03

TOTAL NUMBER OF PAGES IN THIS PUBLICATION IS 522, CONSISTING OF THE FOLLOWING:
Page
No.

* Change
No.

Title . . . . . . . . . . . . . . . . . . . .
T2 . . . . . . . . . . . . . . . . . . . . . .
A .......................
B .......................
C .......................
D .......................
a .......................
b Blank . . . . . . . . . . . . . . . . .
i ........................
ii . . . . . . . . . . . . . . . . . . . . . . .
iii . . . . . . . . . . . . . . . . . . . . . . .
iv . . . . . . . . . . . . . . . . . . . . . . .
v........................
vi . . . . . . . . . . . . . . . . . . . . . . .
vii . . . . . . . . . . . . . . . . . . . . . .
viii . . . . . . . . . . . . . . . . . . . . . .
11 . . . . . . . . . . . . . . . . . . . . . .
12 . . . . . . . . . . . . . . . . . . . . . .
13 . . . . . . . . . . . . . . . . . . . . . .
14 . . . . . . . . . . . . . . . . . . . . . .
15 . . . . . . . . . . . . . . . . . . . . . .
16 . . . . . . . . . . . . . . . . . . . . . .
17 . . . . . . . . . . . . . . . . . . . . . .
18 . . . . . . . . . . . . . . . . . . . . . .
19 . . . . . . . . . . . . . . . . . . . . . .
110 . . . . . . . . . . . . . . . . . . . . .
111 . . . . . . . . . . . . . . . . . . . . .
112 . . . . . . . . . . . . . . . . . . . . .
113 . . . . . . . . . . . . . . . . . . . . .
114 . . . . . . . . . . . . . . . . . . . . .
115 . . . . . . . . . . . . . . . . . . . . .
116 . . . . . . . . . . . . . . . . . . . . .
117 . . . . . . . . . . . . . . . . . . . . .
118 . . . . . . . . . . . . . . . . . . . . .
119 . . . . . . . . . . . . . . . . . . . . .
120 . . . . . . . . . . . . . . . . . . . . .
121 . . . . . . . . . . . . . . . . . . . . .
122 . . . . . . . . . . . . . . . . . . . . .
123 . . . . . . . . . . . . . . . . . . . . .
124 . . . . . . . . . . . . . . . . . . . . .
125 . . . . . . . . . . . . . . . . . . . . .
126 . . . . . . . . . . . . . . . . . . . . .
127 . . . . . . . . . . . . . . . . . . . . .

1
1
1
1
1
1
1
0
1
1
1
0
1
1
1
1
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
1
0
0
1
1
0
0

Page
No.

* ChangeNo.

128 . . . . . . . . . . . . . . . . . . . . .
129 . . . . . . . . . . . . . . . . . . . . .
130 . . . . . . . . . . . . . . . . . . . . .
131 . . . . . . . . . . . . . . . . . . . . .
132 . . . . . . . . . . . . . . . . . . . . .
133 . . . . . . . . . . . . . . . . . . . . .
134 . . . . . . . . . . . . . . . . . . . . .
135 . . . . . . . . . . . . . . . . . . . . .
136 . . . . . . . . . . . . . . . . . . . . .
137 . . . . . . . . . . . . . . . . . . . . .
138 . . . . . . . . . . . . . . . . . . . . .
139 . . . . . . . . . . . . . . . . . . . . .
140 . . . . . . . . . . . . . . . . . . . . .
141 . . . . . . . . . . . . . . . . . . . . .
142 . . . . . . . . . . . . . . . . . . . . .
143 . . . . . . . . . . . . . . . . . . . . .
144 . . . . . . . . . . . . . . . . . . . . .
145 . . . . . . . . . . . . . . . . . . . . .
146 . . . . . . . . . . . . . . . . . . . . .
147 . . . . . . . . . . . . . . . . . . . . .
148 . . . . . . . . . . . . . . . . . . . . .
149 . . . . . . . . . . . . . . . . . . . . .
150 . . . . . . . . . . . . . . . . . . . . .
151 . . . . . . . . . . . . . . . . . . . . .
152 . . . . . . . . . . . . . . . . . . . . .
153 . . . . . . . . . . . . . . . . . . . . .
154 . . . . . . . . . . . . . . . . . . . . .
155 . . . . . . . . . . . . . . . . . . . . .
156 . . . . . . . . . . . . . . . . . . . . .
157 . . . . . . . . . . . . . . . . . . . . .
158 . . . . . . . . . . . . . . . . . . . . .
159 . . . . . . . . . . . . . . . . . . . . .
160 . . . . . . . . . . . . . . . . . . . . .
161 . . . . . . . . . . . . . . . . . . . . .
162 . . . . . . . . . . . . . . . . . . . . .
163 . . . . . . . . . . . . . . . . . . . . .
164 . . . . . . . . . . . . . . . . . . . . .
165 . . . . . . . . . . . . . . . . . . . . .
166 . . . . . . . . . . . . . . . . . . . . .
167 . . . . . . . . . . . . . . . . . . . . .
168 . . . . . . . . . . . . . . . . . . . . .
169 . . . . . . . . . . . . . . . . . . . . .
170 . . . . . . . . . . . . . . . . . . . . .

0
0
0
0
0
0
0
0
0
0
0
0
0
0
1
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
1

Page
No.

* ChangeNo.

171 . . . . . . . . . . . . . . . . . . . . .
172 . . . . . . . . . . . . . . . . . . . . .
173 . . . . . . . . . . . . . . . . . . . . .
174 . . . . . . . . . . . . . . . . . . . . .
175 . . . . . . . . . . . . . . . . . . . . .
176 . . . . . . . . . . . . . . . . . . . . .
177 . . . . . . . . . . . . . . . . . . . . .
178 . . . . . . . . . . . . . . . . . . . . .
179 . . . . . . . . . . . . . . . . . . . . .
180 . . . . . . . . . . . . . . . . . . . . .
181 . . . . . . . . . . . . . . . . . . . . .
182 . . . . . . . . . . . . . . . . . . . . .
183 . . . . . . . . . . . . . . . . . . . . .
184 . . . . . . . . . . . . . . . . . . . . .
185 . . . . . . . . . . . . . . . . . . . . .
186 . . . . . . . . . . . . . . . . . . . . .
187 . . . . . . . . . . . . . . . . . . . . .
188 . . . . . . . . . . . . . . . . . . . . .
189 . . . . . . . . . . . . . . . . . . . . .
190 . . . . . . . . . . . . . . . . . . . . .
191 . . . . . . . . . . . . . . . . . . . . .
192 . . . . . . . . . . . . . . . . . . . . .
193 . . . . . . . . . . . . . . . . . . . . .
194 . . . . . . . . . . . . . . . . . . . . .
195 . . . . . . . . . . . . . . . . . . . . .
196 . . . . . . . . . . . . . . . . . . . . .
197 . . . . . . . . . . . . . . . . . . . . .
198 . . . . . . . . . . . . . . . . . . . . .
199 . . . . . . . . . . . . . . . . . . . . .
1100 . . . . . . . . . . . . . . . . . . . .
1101 . . . . . . . . . . . . . . . . . . . .
1102 . . . . . . . . . . . . . . . . . . . .
1103 . . . . . . . . . . . . . . . . . . . .
1104 . . . . . . . . . . . . . . . . . . . .
1105 . . . . . . . . . . . . . . . . . . . .
1106 . . . . . . . . . . . . . . . . . . . .
1107 . . . . . . . . . . . . . . . . . . . .
1108 . . . . . . . . . . . . . . . . . . . .
1109 . . . . . . . . . . . . . . . . . . . .
1110 . . . . . . . . . . . . . . . . . . . .
1111 . . . . . . . . . . . . . . . . . . . .
1112 . . . . . . . . . . . . . . . . . . . .
1113 . . . . . . . . . . . . . . . . . . . .

0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
1
0
0
0
0
0
0
0
0
0
0
0
0
1
0
0
0

* Zero in this column indicates an original page.

B1
HAF
Change 1A

T.O. GR1F16CJ1INSERT LATEST CHANGED PAGES. DESTROY SUPERSEDED PAGES.

LIST OF EFFECTIVE PAGES (Cont)
Page
No.

*ChangeNo.

1114 . . . . . . . . . . . . . . . . . . . .
1115 . . . . . . . . . . . . . . . . . . . .
1116 . . . . . . . . . . . . . . . . . . . .
1117 . . . . . . . . . . . . . . . . . . . .
1118 . . . . . . . . . . . . . . . . . . . .
1119 . . . . . . . . . . . . . . . . . . . .
1120 . . . . . . . . . . . . . . . . . . . .
1121 . . . . . . . . . . . . . . . . . . . .
1122 . . . . . . . . . . . . . . . . . . . .
1123 . . . . . . . . . . . . . . . . . . . .
1124 . . . . . . . . . . . . . . . . . . . .
1125 . . . . . . . . . . . . . . . . . . . .
1126 . . . . . . . . . . . . . . . . . . . .
1127 . . . . . . . . . . . . . . . . . . . .
1128 . . . . . . . . . . . . . . . . . . . .
1129 . . . . . . . . . . . . . . . . . . . .
1130 . . . . . . . . . . . . . . . . . . . .
1131 . . . . . . . . . . . . . . . . . . . .
1132 . . . . . . . . . . . . . . . . . . . .
1133 . . . . . . . . . . . . . . . . . . . .
1134 . . . . . . . . . . . . . . . . . . . .
1135 . . . . . . . . . . . . . . . . . . . .
1136 . . . . . . . . . . . . . . . . . . . .
1137 . . . . . . . . . . . . . . . . . . . .
1138 . . . . . . . . . . . . . . . . . . . .
1139 . . . . . . . . . . . . . . . . . . . .
1140 . . . . . . . . . . . . . . . . . . . .
1141 . . . . . . . . . . . . . . . . . . . .
1142 . . . . . . . . . . . . . . . . . . . .
1143 . . . . . . . . . . . . . . . . . . . .
1144 . . . . . . . . . . . . . . . . . . . .
1145 . . . . . . . . . . . . . . . . . . . .
1146 . . . . . . . . . . . . . . . . . . . .
1147 . . . . . . . . . . . . . . . . . . . .
1148 . . . . . . . . . . . . . . . . . . . .
1149 . . . . . . . . . . . . . . . . . . . .
1150 . . . . . . . . . . . . . . . . . . . .
1151 . . . . . . . . . . . . . . . . . . . .
1152 . . . . . . . . . . . . . . . . . . . .
1153 . . . . . . . . . . . . . . . . . . . .
1154 . . . . . . . . . . . . . . . . . . . .
1155 . . . . . . . . . . . . . . . . . . . .
1156 . . . . . . . . . . . . . . . . . . . .
1157 . . . . . . . . . . . . . . . . . . . .
1158 . . . . . . . . . . . . . . . . . . . .
1159 . . . . . . . . . . . . . . . . . . . .
1160 . . . . . . . . . . . . . . . . . . . .
1161 . . . . . . . . . . . . . . . . . . . .
1162 . . . . . . . . . . . . . . . . . . . .
1163 . . . . . . . . . . . . . . . . . . . .
1164 . . . . . . . . . . . . . . . . . . . .
1165 . . . . . . . . . . . . . . . . . . . .
1166 . . . . . . . . . . . . . . . . . . . .
1167 . . . . . . . . . . . . . . . . . . . .
1168 . . . . . . . . . . . . . . . . . . . .

0
0
0
0
0
0
0
0
0
0
0
0
0
1
0
0
0
0
0
0
0
0
0
0
0
1
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0

Page
No.

*ChangeNo.

1169 . . . . . . . . . . . . . . . . . . . .
1170 . . . . . . . . . . . . . . . . . . . .
1171 . . . . . . . . . . . . . . . . . . . .
1172 . . . . . . . . . . . . . . . . . . . .
1173 . . . . . . . . . . . . . . . . . . . .
1174 . . . . . . . . . . . . . . . . . . . .
1175 . . . . . . . . . . . . . . . . . . . .
1176 . . . . . . . . . . . . . . . . . . . .
1177 . . . . . . . . . . . . . . . . . . . .
1178 . . . . . . . . . . . . . . . . . . . .
1179 . . . . . . . . . . . . . . . . . . . .
1180 . . . . . . . . . . . . . . . . . . . .
1181 . . . . . . . . . . . . . . . . . . . .
1182 . . . . . . . . . . . . . . . . . . . .
1183 . . . . . . . . . . . . . . . . . . . .
1184 . . . . . . . . . . . . . . . . . . . .
1185 . . . . . . . . . . . . . . . . . . . .
1186 . . . . . . . . . . . . . . . . . . . .
1187 . . . . . . . . . . . . . . . . . . . .
1188 . . . . . . . . . . . . . . . . . . . .
1189 . . . . . . . . . . . . . . . . . . . .
1190 . . . . . . . . . . . . . . . . . . . .
1191 . . . . . . . . . . . . . . . . . . . .
1192 . . . . . . . . . . . . . . . . . . . .
1193 . . . . . . . . . . . . . . . . . . . .
1194 . . . . . . . . . . . . . . . . . . . .
1195 . . . . . . . . . . . . . . . . . . . .
1196 . . . . . . . . . . . . . . . . . . . .
1197 . . . . . . . . . . . . . . . . . . . .
1198 . . . . . . . . . . . . . . . . . . . .
1199 . . . . . . . . . . . . . . . . . . . .
1200 . . . . . . . . . . . . . . . . . . . .
1201 . . . . . . . . . . . . . . . . . . . .
1202 . . . . . . . . . . . . . . . . . . . .
1203 . . . . . . . . . . . . . . . . . . . .
1204 . . . . . . . . . . . . . . . . . . . .
1205 . . . . . . . . . . . . . . . . . . . .
1206 . . . . . . . . . . . . . . . . . . . .
1207 . . . . . . . . . . . . . . . . . . . .
1208 . . . . . . . . . . . . . . . . . . . .
1209 . . . . . . . . . . . . . . . . . . . .
1210 . . . . . . . . . . . . . . . . . . . .
1211 . . . . . . . . . . . . . . . . . . . .
1212 . . . . . . . . . . . . . . . . . . . .
1213 . . . . . . . . . . . . . . . . . . . .
1214 . . . . . . . . . . . . . . . . . . . .
1215 . . . . . . . . . . . . . . . . . . . .
1216 . . . . . . . . . . . . . . . . . . . .
1217 . . . . . . . . . . . . . . . . . . . .
1218 . . . . . . . . . . . . . . . . . . . .
1219 . . . . . . . . . . . . . . . . . . . .
1220 . . . . . . . . . . . . . . . . . . . .
1221 . . . . . . . . . . . . . . . . . . . .
1222 . . . . . . . . . . . . . . . . . . . .
1223 . . . . . . . . . . . . . . . . . . . .

0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
1
1
1
0
1
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
1

* Zero in this column indicates an original page.

BChange 1

Page
No.

*ChangeNo.

1224 . . . . . . . . . . . . . . . . . . . .
1225 . . . . . . . . . . . . . . . . . . . .
1226 . . . . . . . . . . . . . . . . . . . .
21 . . . . . . . . . . . . . . . . . . . . . .
22 . . . . . . . . . . . . . . . . . . . . . .
23 . . . . . . . . . . . . . . . . . . . . . .
24 . . . . . . . . . . . . . . . . . . . . . .
25 . . . . . . . . . . . . . . . . . . . . . .
26 . . . . . . . . . . . . . . . . . . . . . .
27 . . . . . . . . . . . . . . . . . . . . . .
28 . . . . . . . . . . . . . . . . . . . . . .
29 . . . . . . . . . . . . . . . . . . . . . .
210 . . . . . . . . . . . . . . . . . . . . .
211 . . . . . . . . . . . . . . . . . . . . .
212 . . . . . . . . . . . . . . . . . . . . .
213 . . . . . . . . . . . . . . . . . . . . .
214 . . . . . . . . . . . . . . . . . . . . .
215 . . . . . . . . . . . . . . . . . . . . .
216 . . . . . . . . . . . . . . . . . . . . .
217 . . . . . . . . . . . . . . . . . . . . .
218 . . . . . . . . . . . . . . . . . . . . .
219 . . . . . . . . . . . . . . . . . . . . .
220 . . . . . . . . . . . . . . . . . . . . .
221 . . . . . . . . . . . . . . . . . . . . .
222 . . . . . . . . . . . . . . . . . . . . .
223 . . . . . . . . . . . . . . . . . . . . .
224 . . . . . . . . . . . . . . . . . . . . .
225 . . . . . . . . . . . . . . . . . . . . .
226 . . . . . . . . . . . . . . . . . . . . .
227 . . . . . . . . . . . . . . . . . . . . .
228 . . . . . . . . . . . . . . . . . . . . .
229 . . . . . . . . . . . . . . . . . . . . .
230 . . . . . . . . . . . . . . . . . . . . .
231 . . . . . . . . . . . . . . . . . . . . .
232 . . . . . . . . . . . . . . . . . . . . .
233 . . . . . . . . . . . . . . . . . . . . .
234 . . . . . . . . . . . . . . . . . . . . .
235 . . . . . . . . . . . . . . . . . . . . .
236 . . . . . . . . . . . . . . . . . . . . .
237 . . . . . . . . . . . . . . . . . . . . .
238 . . . . . . . . . . . . . . . . . . . . .
239 . . . . . . . . . . . . . . . . . . . . .
240 . . . . . . . . . . . . . . . . . . . . .
241 . . . . . . . . . . . . . . . . . . . . .
242 . . . . . . . . . . . . . . . . . . . . .
31 . . . . . . . . . . . . . . . . . . . . . .
32 . . . . . . . . . . . . . . . . . . . . . .
33 . . . . . . . . . . . . . . . . . . . . . .
34 . . . . . . . . . . . . . . . . . . . . . .
35 . . . . . . . . . . . . . . . . . . . . . .
36 . . . . . . . . . . . . . . . . . . . . . .
37 . . . . . . . . . . . . . . . . . . . . . .
38 . . . . . . . . . . . . . . . . . . . . . .
39 . . . . . . . . . . . . . . . . . . . . . .
310 . . . . . . . . . . . . . . . . . . . . .

0
0
0
0
0
0
0
1
0
0
0
0
0
0
1
1
0
0
0
1
0
0
0
0
0
1
0
0
1
0
0
0
1
0
0
0
1
1
0
0
1
0
0
0
0
0
0
0
0
0
0
0
0
1
0

T.O. GR1F16CJ1INSERT LATEST CHANGED PAGES. DESTROY SUPERSEDED PAGES.

LIST OF EFFECTIVE PAGES (Cont)
Page
No.

*ChangeNo.

311 . . . . . . . . . . . . . . . . . . . . .
312 . . . . . . . . . . . . . . . . . . . . .
313 . . . . . . . . . . . . . . . . . . . . .
314 . . . . . . . . . . . . . . . . . . . . .
315 . . . . . . . . . . . . . . . . . . . . .
316 . . . . . . . . . . . . . . . . . . . . .
317 . . . . . . . . . . . . . . . . . . . . .
318 . . . . . . . . . . . . . . . . . . . . .
319 . . . . . . . . . . . . . . . . . . . . .
320 . . . . . . . . . . . . . . . . . . . . .
321 . . . . . . . . . . . . . . . . . . . . .
322 . . . . . . . . . . . . . . . . . . . . .
323 . . . . . . . . . . . . . . . . . . . . .
324 . . . . . . . . . . . . . . . . . . . . .
325 . . . . . . . . . . . . . . . . . . . . .
326 . . . . . . . . . . . . . . . . . . . . .
327 . . . . . . . . . . . . . . . . . . . . .
328 . . . . . . . . . . . . . . . . . . . . .
329 . . . . . . . . . . . . . . . . . . . . .
330 . . . . . . . . . . . . . . . . . . . . .
331 . . . . . . . . . . . . . . . . . . . . .
332 . . . . . . . . . . . . . . . . . . . . .
333 . . . . . . . . . . . . . . . . . . . . .
334 . . . . . . . . . . . . . . . . . . . . .
335 . . . . . . . . . . . . . . . . . . . . .
336 . . . . . . . . . . . . . . . . . . . . .
337 . . . . . . . . . . . . . . . . . . . . .
338 . . . . . . . . . . . . . . . . . . . . .
339 . . . . . . . . . . . . . . . . . . . . .
340 . . . . . . . . . . . . . . . . . . . . .
340.1 Added . . . . . . . . . . . . .
340.2 Blank Added . . . . . .
341 . . . . . . . . . . . . . . . . . . . . .
342 . . . . . . . . . . . . . . . . . . . . .
343 . . . . . . . . . . . . . . . . . . . . .
344 . . . . . . . . . . . . . . . . . . . . .
345 . . . . . . . . . . . . . . . . . . . . .
346 . . . . . . . . . . . . . . . . . . . . .
347 . . . . . . . . . . . . . . . . . . . . .
348 . . . . . . . . . . . . . . . . . . . . .
349 . . . . . . . . . . . . . . . . . . . . .
350 . . . . . . . . . . . . . . . . . . . . .
351 . . . . . . . . . . . . . . . . . . . . .
352 . . . . . . . . . . . . . . . . . . . . .
353 . . . . . . . . . . . . . . . . . . . . .
354 . . . . . . . . . . . . . . . . . . . . .
355 . . . . . . . . . . . . . . . . . . . . .
356 . . . . . . . . . . . . . . . . . . . . .
357 . . . . . . . . . . . . . . . . . . . . .
358 . . . . . . . . . . . . . . . . . . . . .
359 . . . . . . . . . . . . . . . . . . . . .
360 . . . . . . . . . . . . . . . . . . . . .
361 . . . . . . . . . . . . . . . . . . . . .
362 . . . . . . . . . . . . . . . . . . . . .
363 . . . . . . . . . . . . . . . . . . . . .

0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
1
0
0
0
0
0
0
0
0
0
0
1
1
0
1
1
1
1
0
0
1
0
0
0
0
1
1
0
0
0
0
0
0
0
0
0
1
0
0
0

Page
No.

*ChangeNo.

364 . . . . . . . . . . . . . . . . . . . . .
365 . . . . . . . . . . . . . . . . . . . . .
366 . . . . . . . . . . . . . . . . . . . . .
367 . . . . . . . . . . . . . . . . . . . . .
368 . . . . . . . . . . . . . . . . . . . . .
369 . . . . . . . . . . . . . . . . . . . . .
370 . . . . . . . . . . . . . . . . . . . . .
371 . . . . . . . . . . . . . . . . . . . . .
372 . . . . . . . . . . . . . . . . . . . . .
373 . . . . . . . . . . . . . . . . . . . . .
374 . . . . . . . . . . . . . . . . . . . . .
375 . . . . . . . . . . . . . . . . . . . . .
376 . . . . . . . . . . . . . . . . . . . . .
377 . . . . . . . . . . . . . . . . . . . . .
378 . . . . . . . . . . . . . . . . . . . . .
379 . . . . . . . . . . . . . . . . . . . . .
380 . . . . . . . . . . . . . . . . . . . . .
381 . . . . . . . . . . . . . . . . . . . . .
382 . . . . . . . . . . . . . . . . . . . . .
383 . . . . . . . . . . . . . . . . . . . . .
384 . . . . . . . . . . . . . . . . . . . . .
384.1 Added . . . . . . . . . . . . .
384.2 Blank Added . . . . . .
385 . . . . . . . . . . . . . . . . . . . . .
386 . . . . . . . . . . . . . . . . . . . . .
387 . . . . . . . . . . . . . . . . . . . . .
388 . . . . . . . . . . . . . . . . . . . . .
389 . . . . . . . . . . . . . . . . . . . . .
390 . . . . . . . . . . . . . . . . . . . . .
391 . . . . . . . . . . . . . . . . . . . . .
392 . . . . . . . . . . . . . . . . . . . . .
393 . . . . . . . . . . . . . . . . . . . . .
394 . . . . . . . . . . . . . . . . . . . . .
395 . . . . . . . . . . . . . . . . . . . . .
396 . . . . . . . . . . . . . . . . . . . . .
397 . . . . . . . . . . . . . . . . . . . . .
398 . . . . . . . . . . . . . . . . . . . . .
399 . . . . . . . . . . . . . . . . . . . . .
3100 . . . . . . . . . . . . . . . . . . . .
3101 . . . . . . . . . . . . . . . . . . . .
3102 . . . . . . . . . . . . . . . . . . . .
3103 . . . . . . . . . . . . . . . . . . . .
3104 . . . . . . . . . . . . . . . . . . . .
3105 . . . . . . . . . . . . . . . . . . . .
3106 . . . . . . . . . . . . . . . . . . . .
3107 . . . . . . . . . . . . . . . . . . . .
3108 . . . . . . . . . . . . . . . . . . . .
3109 . . . . . . . . . . . . . . . . . . . .
3110 . . . . . . . . . . . . . . . . . . . .
3111 . . . . . . . . . . . . . . . . . . . .
3112 . . . . . . . . . . . . . . . . . . . .
3113 . . . . . . . . . . . . . . . . . . . .
3114 . . . . . . . . . . . . . . . . . . . .
3115 . . . . . . . . . . . . . . . . . . . .
3116 . . . . . . . . . . . . . . . . . . . .

1
0
1
1
0
0
0
1
0
0
0
0
0
0
0
0
1
1
1
0
1
1
1
1
0
1
0
0
0
0
0
0
0
0
1
1
1
1
0
1
1
1
0
1
1
1
0
0
0
1
1
1
0
0
0

Page
No.

*ChangeNo.

3117 . . . . . . . . . . . . . . . . . . . .
3118 . . . . . . . . . . . . . . . . . . . .
3119 . . . . . . . . . . . . . . . . . . . .
3120 . . . . . . . . . . . . . . . . . . . .
3121 . . . . . . . . . . . . . . . . . . . .
3122 . . . . . . . . . . . . . . . . . . . .
3122.1 Added . . . . . . . . . . .
3122.2 Blank Added . . . . .
3123 . . . . . . . . . . . . . . . . . . . .
3124 . . . . . . . . . . . . . . . . . . . .
3125 . . . . . . . . . . . . . . . . . . . .
3126 . . . . . . . . . . . . . . . . . . . .
3127 . . . . . . . . . . . . . . . . . . . .
3128 . . . . . . . . . . . . . . . . . . . .
3129 . . . . . . . . . . . . . . . . . . . .
3130 . . . . . . . . . . . . . . . . . . . .
3131 . . . . . . . . . . . . . . . . . . . .
3132 . . . . . . . . . . . . . . . . . . . .
3133 . . . . . . . . . . . . . . . . . . . .
3134 . . . . . . . . . . . . . . . . . . . .
3135 . . . . . . . . . . . . . . . . . . . .
3136 . . . . . . . . . . . . . . . . . . . .
3137 . . . . . . . . . . . . . . . . . . . .
3138 . . . . . . . . . . . . . . . . . . . .
3139 . . . . . . . . . . . . . . . . . . . .
3140 . . . . . . . . . . . . . . . . . . . .
3141 . . . . . . . . . . . . . . . . . . . .
3142 . . . . . . . . . . . . . . . . . . . .
3143 . . . . . . . . . . . . . . . . . . . .
3144 . . . . . . . . . . . . . . . . . . . .
3145 . . . . . . . . . . . . . . . . . . . .
3146 Blank . . . . . . . . . . . . .
41 . . . . . . . . . . . . . . . . . . . . . .
42 Blank . . . . . . . . . . . . . . . .
51 . . . . . . . . . . . . . . . . . . . . . .
52 . . . . . . . . . . . . . . . . . . . . . .
53 . . . . . . . . . . . . . . . . . . . . . .
54 . . . . . . . . . . . . . . . . . . . . . .
55 . . . . . . . . . . . . . . . . . . . . . .
56 . . . . . . . . . . . . . . . . . . . . . .
57 . . . . . . . . . . . . . . . . . . . . . .
58 . . . . . . . . . . . . . . . . . . . . . .
59 . . . . . . . . . . . . . . . . . . . . . .
510 . . . . . . . . . . . . . . . . . . . . .
511 . . . . . . . . . . . . . . . . . . . . .
512 . . . . . . . . . . . . . . . . . . . . .
513 . . . . . . . . . . . . . . . . . . . . .
514 . . . . . . . . . . . . . . . . . . . . .
515 . . . . . . . . . . . . . . . . . . . . .
516 Blank . . . . . . . . . . . . . . .
61 . . . . . . . . . . . . . . . . . . . . . .
62 . . . . . . . . . . . . . . . . . . . . . .
63 . . . . . . . . . . . . . . . . . . . . . .
64 . . . . . . . . . . . . . . . . . . . . . .
65 . . . . . . . . . . . . . . . . . . . . . .

0
0
1
0
0
1
1
1
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
1
0
0
0
0
0
0
0
0
0
0
1
0
1
0
0
0
0
0
0

* Zero in this column indicates an original page.

Change 1C

T.O. GR1F16CJ1INSERT LATEST CHANGED PAGES. DESTROY SUPERSEDED PAGES.

LIST OF EFFECTIVE PAGES (Cont)
Page
No.

*ChangeNo.

66 . . . . . . . . . . . . . . . . . . . . . .
67 . . . . . . . . . . . . . . . . . . . . . .
68 . . . . . . . . . . . . . . . . . . . . . .
69 . . . . . . . . . . . . . . . . . . . . . .
610 . . . . . . . . . . . . . . . . . . . . .
611 . . . . . . . . . . . . . . . . . . . . .
612 . . . . . . . . . . . . . . . . . . . . .
613 . . . . . . . . . . . . . . . . . . . . .
614 . . . . . . . . . . . . . . . . . . . . .
615 . . . . . . . . . . . . . . . . . . . . .
616 . . . . . . . . . . . . . . . . . . . . .
71 . . . . . . . . . . . . . . . . . . . . . .
72 . . . . . . . . . . . . . . . . . . . . . .
73 . . . . . . . . . . . . . . . . . . . . . .
74 . . . . . . . . . . . . . . . . . . . . . .
75 . . . . . . . . . . . . . . . . . . . . . .
76 . . . . . . . . . . . . . . . . . . . . . .
77 . . . . . . . . . . . . . . . . . . . . . .
78 . . . . . . . . . . . . . . . . . . . . . .
81 Added . . . . . . . . . . . . . . .
82 Added . . . . . . . . . . . . . . .
83 Added . . . . . . . . . . . . . . .
84 Added . . . . . . . . . . . . . . .
85 Added . . . . . . . . . . . . . . .
86 Added . . . . . . . . . . . . . . .
87 Added . . . . . . . . . . . . . . .
88 Added . . . . . . . . . . . . . . .
89 Added . . . . . . . . . . . . . . .
810 Added . . . . . . . . . . . . . .
811 Added . . . . . . . . . . . . . .
812 Added . . . . . . . . . . . . . .
813 Added . . . . . . . . . . . . . .
814 Added . . . . . . . . . . . . . .
815 Added . . . . . . . . . . . . . .
816 Added . . . . . . . . . . . . . .
817 Added . . . . . . . . . . . . . .
818 Added . . . . . . . . . . . . . .
819 Added . . . . . . . . . . . . . .
820 Added . . . . . . . . . . . . . .
821 Added . . . . . . . . . . . . . .
822 Added . . . . . . . . . . . . . .
823 Added . . . . . . . . . . . . . .
824 Added . . . . . . . . . . . . . .
825 Added . . . . . . . . . . . . . .
826 Added . . . . . . . . . . . . . .
827 Added . . . . . . . . . . . . . .
828 Added . . . . . . . . . . . . . .
829 Added . . . . . . . . . . . . . .
830 Added . . . . . . . . . . . . . .
831 Added . . . . . . . . . . . . . .
832 Added . . . . . . . . . . . . . .
833 Added . . . . . . . . . . . . . .
834 Added . . . . . . . . . . . . . .
835 Added . . . . . . . . . . . . . .
836 Added . . . . . . . . . . . . . .

0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1

Page
No.

*ChangeNo.

837 Added . . . . . . . . . . . . . .
838 Blank Added . . . . . . . .
Glossary 1 . . . . . . . . . . . . . . .
Glossary 2 . . . . . . . . . . . . . . .
Glossary 3 . . . . . . . . . . . . . . .
Glossary 4 . . . . . . . . . . . . . . .
Glossary 5 . . . . . . . . . . . . . . .
Glossary 6 Blank . . . . . . . . .

1
1
1
1
1
1
0
0

* Zero in this column indicates an original page.

DChange 1

Page
No.

*ChangeNo.

T.O. GR1F16CJ1FLIGHT MANUALS, SAFETY SUPPLEMENT, AND
OPERATIONAL SUPPLEMENT STATUS
This supplement status page is based on information available as of 15 June 2003. It is not an official status page.
Flight Manual

Basic Date

Change No. and Date

T.O. GR1F16CJ1

15 Oct 2002

1

Supplemental Flight Manuals

Basic Date

Change No. and Date

T.O. GR1F16CJ11

9 Jun 19977

15 Nov 2002

T.O. GR1F16CJ12

15 Feb 2000

515 Mar 2003

Flight Crew Checklist

Basic Date

Change No. and Date

T.O. GR1F16CJ1CL1

15 Oct 2002

1

15 Jun 2003

15 Jun 2003

INCORPORATED SAFETY AND OPERATIONAL SUPPLEMENTS
Flight Manual
Sections Affected

T.O. Number

Date

Short Title

GR1F16CJ1SS27

27 Nov 02

FLCS Failure That Results
In Loss Of Pitch Control

III

GR1F16CJ1SS29

29 Jan 03

Max Range Speeds and FO Pattern
Altitudes

III

GR1F16CJ1S31

30 Apr 03

Fuel Transfer Switch, Avionics
Fault & ECM Lights, and EGI
Align

II

OUTSTANDING SAFETY AND OPERATIONAL SUPPLEMENTS
T.O. Number

Date

Short Title

Flight Manual
Sections Affected

None

Change 1a/(b blank)

T.O. GR1F16CJ1TABLE OF CONTENTSSECTION I

Description and Operation

11

SECTION II

Normal Procedures

21

SECTION III

Emergency Procedures

31

SECTION IV

Crew Duties

41

SECTION V

Operating Limitations

51

SECTION V

Operating Limitations (Stores)

SECTION VI

Flight Characteristics

61

SECTION VII

Adverse Weather Operation

71

SECTION VIII

Air Refueling Procedures

81

GLOSSARYStandard and Nonstandard Abbreviations

APPENDIX A

Performance Data-F100PW229

T.O. GR1F16CJ11

APPENDIX B

Performance Data-F110GE129

T.O. GR1F16CJ11

APPENDIX C

Performance Data-F100PW229/CFT

T.O. GR1F16CJ11

T.O. GR1F16CJ12

Glossary 1

Change 1i

T.O. GR1F16CJ1

BEFOREYOU
TAKE OFF,READ THIS!

1F-16X-1-0002X

SCOPE
This manual contains the necessary information forsafe and efficient operation of the aircraft. These
instructions provide a general knowledge of theaircraft and its characteristics and specific normal
and emergency operating procedures. Pilot experi
ence is recognized; therefore, basic flight principles
are avoided. Instructions in this manual are prepared
to be understandable to the least experienced pilot
who can be expected to operate the aircraft. Thismanual provides the best possible operating instruc
tions under most conditions. Multiple emergencies,
adverse weather, terrain, etc., may require modifica
tion of the procedures. This manual must be usedwith one or more of the following manuals to obtain
information necessary for safe and efficient opera
tion:
T.O. GR1F16CJ11

Supplemental Flight
Manual, F16C/D Air
craft

T.O. GR1F16CJ12

Supplemental Flight
Manual, F16C/D Air
craft

T.O. GR1F16CJ52

Loading DataiiChange 1

T.O. GR1F16CJ6CF1

Acceptance and Func
tional Check Flight
Procedures Manual,F16C/D Aircraft

T.O. GR1F16CJ3411

Avionics and Nonnu
clear Weapons Deliv
ery Flight Manual

T.O. GR1F16CJ3412

Nonnuclear Weapons
Ballistics

T.O. 11C1

Basic Flight Crew Air
Refueling Procedures

PERMISSIBLE OPERATIONS
The flight manual takes a positive approach andnormally states only what can be done. Unusual
operations or configurations are prohibited unlessspecifically covered herein. Clearance must be
obtained before any questionable operation which is
not specifically permitted in this manual is
attempted.

T.O. GR1F16CJ1HOW TO BE ASSURED OF HAVING LATEST DATASAFETY SUPPLEMENTS

Refer to T.O. GR0171 for a listing of all current flight
manuals, safety supplements, operational supple
ments, and checklists. Also, check the flight manual
title page, the title block of each safety andoperational supplement, and all status pages
contained in the flight manual or attached to formal
safety and operational supplements. Clear all
discrepancies before flight.

Information involving safety is promptly forwarded
in a safety supplement. Urgent information is
published in interim safety supplements andtransmitted by electronic message. Formal supple
ments are mailed. The supplement title block andstatus page (published with formal supplementsonly) should be checked to determine the effect of the
supplement on this manual and other outstanding
supplements.

ARRANGEMENT
The manual is divided into eight sections and three
supplemental appendices.
ILLUSTRATIONS
Cockpit arrangement, cockpit console, and cockpit
instrument panel illustrations display the delivered
configuration plus the approved equipment modifica
tions. For details of equipment modification, see theindividual equipment illustration.
SUPPLEMENT INFORMATION AND GUIDELINES
Supplements are safety or operational and are
indicated 1SS or 1S, respectively. Supplements areissued as interim electronic messages or formal
printed copies. All interim supplements are assigned
odd numbers, such as 1SS195. When an interim
supplement is formalized, it is assigned the next
following even number, such as 1SS196. Formal
supplements not preceded by an interim supplement
are also assigned even numbers. If an interim
supplement is not to be formalized, a statement
cancelling the next assigned even supplement
number is included in the REMARKS section of theinterim supplement. If a formal supplement is notpreceded by an interim supplement, a statement
cancelling the previous odd supplement number is
included on the status page of the formal supplement.
Occasionally, a supplement has dual references in the
instructions; this is because the supplement applies
to the present and subsequent manual. Minor
text/illustration changes or deletions are given as
instructions in the supplement. When lengthy
additions are required, the formal supplement
provides onesided insert page(s) to the flight manualand checklist. This supplement page(s) is attached tothe original page(s). The original page(s) remains in
the manual or checklist in case the supplement is
rescinded and the page(s) is needed. Added page(s)(e.g., 348.1) are inserted in proper numericalsequence and may be printed on both sides.

OPERATIONAL SUPPLEMENTS
Information involving changes to operating procedures is forwarded by operational supplements. The
procedure for handling operational supplements isthe same as for safety supplements.
CHECKLISTThe checklist contains itemized procedures without
all of the amplification. Primary line items in theflight manual and checklist are identical.
HOW TO GET PERSONAL COPIESEach pilot is entitled to a personal copy of the flight
manual, safety supplements, operational supple
ments, and a checklist. The required quantities
should be ordered before needed to assure theirtimely receipt. Check with the publication distribu
tion officer whose job is to fulfill T.O. requests.
Country manuals and difference data are assigned aCountry Standard T.O. (CSTO) number and indexedin the Security Assistance Program T.O. Index, T.O.
GR0171. Updates, stock, store, and issue of CSTO's
are a contractor supported function. Insure a systemis established at each base to deliver the publications
to the pilots immediately upon receipt.
FLIGHT MANUAL BINDERS
Looseleaf binders and sectionalized tabs are avail
able for use with the manual. They are obtained
through local purchase procedures and are listed in
the USAF supply channels and may be ordered on the
supply case through AFLC. Check with supply
personnel for assistance in procuring these items.
CHANGE SYMBOLThe change symbol, as illustrated by the black line in
the margin of this paragraph, indicates changes
made to the current issue.

Change 1iii

T.O. GR1F16CJ1WARNINGS, CAUTIONS, AND NOTESThe following definitions apply to Warnings,
Cautions, and Notes found throughout the manual:

regarding this manual or any phase of the flight
manual program are welcomed. These should be
forwarded on AF Form 847 in accordance with AFI
11215 through command headquarters to OOALC/
YPVT, 6080 Gum Lane, Hill AFB, UT 840565825.
PUBLICATION DATE

Operating procedures, techniques,etc., which could result in personal
injury or loss of life if not carefully
followed.

Operating procedures, techniques,etc., which could result in damage toequipment if not carefully followed.

The date appearing on the title page represents thecurrency of material contained herein.
AIRCRAFT AND COCKPIT DESIGNATION CODES
System and/or component effectivity for a particularaircraft version/cockpit is denoted by a letter code
enclosed in a box located in the text or on an
illustration. The symbols and designations are as
follows:
No code-F16C and F16D aircraft

NOTE

C

F16C aircraft

An operating procedure, technique,etc., which is considered essential toemphasize.

D

F16D aircraft

DF

F16D aircraft, forward cockpit

DRF16D aircraft, rear cockpit

PX II

Peace Xenia II aircraft

PX III

Peace Xenia III aircraft

USE OF WORDS SHALL, WILL, SHOULD, AND
MAY
The word shall or will is used to indicate a mandatory
requirement. The word should is used to indicate anonmandatory desired or preferred method ofaccomplishment. The word may is used to indicate an
acceptable or suggested means of accomplishment.
USE OF WORDS AS DESIRED AND AS REQUIREDAs desired allows pilot preference in switch/control
positioning.
As required indicates those actions which vary basedon mission requirements.

ENGINE DESIGNATION CODES
System and/or component effectivity for a particularengine version is denoted by an engine code enclosed
in a box located in the text or on an illustration. The
symbols and designations are as follows:
No code-Either engine
PW 229F100PW229

GE 129F110GE129

AIRSPEED REFERENCES
All references to airspeed quoted in knots refer toindicated airspeed.
PILOT'S RESPONSIBILITY-TO LET US KNOW
Every effort is made to keep the flight manualcurrent. Review conferences with operating person
nel and a constant review of safety investigation andflight test reports assure inclusion of the latest datain the manual. Comments, corrections, and questionsivBLOCK DESIGNATION CODES/SERIAL NUMBER/
TAIL NUMBER CROSSREFERENCE
Because of differences in configuration between
aircraft and to avoid repetitious use of aircraft serial
numbers, a block effectivity system is used. The block
effectivities reflect the aircraft block, the aircraft
serial number, and the tail number. Attrited aircraft
are removed from the listing. This system is usedthroughout the manual, both in text and illustra
tions.

T.O. GR1F16CJ1AIRCRAFT MODIFICATION/RETROFIT INFORMA
TION
This list includes the applicable T.O./ECP effectivi
ties. It is not an official status page. Refer to T.O. 014
for the complete listing of TCTO's. Throughout thismanual, black TV screen symbols containing white
numerals ((2)) are used to distinguish information
related to aircraft which are modified by a specific
TV CODE/
T.O. NO.
b1q

1F162121

SHORT TITLE
Flight Control Com
puter Software Update (ECP 2269)

T.O./ECP. Information pertaining to modified aircraft
is identified by an appropriate effectivity symbol.
Information which is not identified by an effectivity
symbol is considered common to all aircraft.
Information pertaining only to unmodified aircraft
has the appropriate effectivity symbol preceded by
LESS. For example, LESS (4) indicates that the
information is only applicable to aircraft not modified
by a specific T.O./ECP.
EFFECTIVITYPRODUCTIONRETROFITF16C 931055,
and onF16C 931045 thru
931054

F16D NAF16D All

Summary:DD

b2p

Retrofit authorized
Modification incorporates FLCS improvements, including improved reset capability for FLCC power up failures and elimination of lateral accelerometer failures during gun firing.

1F162170

Modification of theEjection Mode Selec
tor Valve Console(ECP 2362)

F16C NAF16C NAF16D 991534,
and onF16D 931077 thru
931084

F16C 991500,
and onF16C 931045 thru
931076

F16D 991534,
and onF16D 931077 thru
931084

Summary:
D Retrofit authorized
D Modification enlarges the hole in the console to allow the EJECTION MODE SEL handle to seatproperly in the NORM position.
b2t

XX1F162077

Summary:
D
D

b6n

NA

Installation of
Improved Antiskid
Braking System(OCP H086)(ECP 2478)

Retrofit authorized
Modification incorporates an improved Brake Control/Antiskid Assembly to replace the current
Brake Control and Antiskid Control Boxes.

Flight Control Com
puter Software Update for CFT(ECP 2495)

F16C 991500,
and onNA

F16D 991534,
Summary:
and on
D Retrofit NA
D Modification incorporates flight control system software updates to provide increased roll rate capa
bility in CAT III with CFT's and eliminate an existing prohibited maneuver.

Change 1v

T.O. GR1F16CJ1TV CODE/
T.O. NO.
b7n

TBD

SHORT TITLE
AVIONICS FAULT
caution/ECM lights
CDEEU mod
(ECP 2526)

EFFECTIVITYPRODUCTIONRETROFITF16C 991507,
and onTBD

F16D 991541,
Summary:
and on
D Retrofit TBD
D Modification to circuitry in the CDEEU which will allow the AVIONICS FAULT caution and ECM
lights to illuminate in both cockpits of PX III D aircraft.

viChange 1

T.O. GR1F16CJ1Block Designation Codes/Serial
Number/Tail Number CrossĆReference
BLOCK 50
Serial Numbers/
Tail Numbers

C
PX II
931045/045
931046/046
931047/047
931048/048
931049/049
931050/050
931051/051
931052/052
931053/053
931054/054
931055/055
931056/056
931057/057
931058/058
931059/059
931060/060
931061/061
931062/062
931063/063
931064/064
931065/065
931066/066
931067/067
931068/068
931069/069
931070/070
931071/071
931072/072
931073/073
931074/074
931075/075
931076/076

D
PX II
931077/077
931078/078
931079/079

931080/080
931081/081
931082/082
931083/083
931084/084

BLOCK 52+
Serial Numbers/
Tail NumbersC
PX III991500/500
991501/501
991502/502
991503/503
991504/504
991505/505
991506/506
991507/507
991508/508
991509/509
991510/510
991511/511
991512/512
991513/513
991514/514
991515/515
991516/516
991517/517
991518/518
991519/519
991520/520
991521/521
991522/522
991523/523
991524/524
991525/525
991526/526
991527/527
991528/528
991529/529
991530/530
991531/531
991532/532

991533/533
018530/534
018531/535
018532/536
018533/537
018534/538
018535/539

D
PX III991534/600
991535/601
991536/602
991537/603
991538/604
991539/605
991540/606
991541/607
991542/608
991543/609
991544/610
991545/611
991546/612
991547/613
991548/614
991549/615
018536/616
018537/617
018538/618
018539/619

Change 1vii

T.O. GR1F16CJ1C

D

GR1F-16CJ-1-0003A37

viiiChange 1

T.O. GR1F16CJ1

SECTION I
DESCRIPTION AND OPERATIONTABLE OF CONTENTS
The Aircraft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Aircraft General Data . . . . . . . . . . . . . . . . . . .
Aircraft General Arrangement . . . . . . . . . . .
Aircraft Gross Weight (A/C GW) . . . . . . . . . .
Cockpit Arrangement . . . . . . . . . . . . . . . . . . . .
Engine PW 229 . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Engine General Description PW 229 . . . . . .
Engine Fuel System PW 229 . . . . . . . . . . . . .
Engine Control System PW 229 . . . . . . . . . . .
Main Fuel Control (MFC) PW 229 . . . . . .
Afterburner (AB) FuelControl PW 229 . . . . . . . . . . . . . . . . . . . .
Digital Electronic Engine Control(DEEC) PW 229 . . . . . . . . . . . . . . . . . . . .
Secondary Engine Control(SEC) PW 229 . . . . . . . . . . . . . . . . . . . . . .
Main Fuel Pump PW 229 . . . . . . . . . . . . . .
Afterburner (AB) Fuel
Pump PW 229 . . . . . . . . . . . . . . . . . . . . . .
Compressor Inlet Variable Vane
(CIVV) Control PW 229 . . . . . . . . . . . . .
Rear Compressor Variable Vanes(RCVV's) PW 229 . . . . . . . . . . . . . . . . . . .
Compressor Bleed Air PW 229 . . . . . . . . .
Pressurization and Dump
Valve PW 229 . . . . . . . . . . . . . . . . . . . . . .
Exhaust Nozzle PW 229 . . . . . . . . . . . . . . . . . .
Convergent Exhaust Nozzle Control(CENC) PW 229 . . . . . . . . . . . . . . . . . . . .
LightOff Detector (LOD) PW 229 . . . . . .
Engine Diagnostic Unit (EDU) PW 229 . . . .
Engine Oil System PW 229 . . . . . . . . . . . . . . .
FUEL/OIL HOT Caution
Light PW 229 . . . . . . . . . . . . . . . . . . . . . .
Engine AntiIce System PW 229 . . . . . . . . . .
Engine ANTI ICE Switch PW 229 . . . . . .
INLET ICING Caution
Light PW 229 . . . . . . . . . . . . . . . . . . . . . .
Engine and Accessory Drive
Gearboxes PW 229 . . . . . . . . . . . . . . . . . . . .
Engine Alternator PW 229 . . . . . . . . . . . . . . .
Engine Ignition System PW 229 . . . . . . . . . .

16
17
17
17
17
131
131
131
131
131
131
133
133
133
133
133
133
134
134
134
134
134
134
134
134
135
135
135
135
135
135

Jet Fuel Starter (JFS) PW 229 . . . . . . . . . . . .
ENG & JET START Control
Panel PW 229 . . . . . . . . . . . . . . . . . . . . . . . .
JFS Switch C DF PW 229 . . . . . . . . . . . . .
JFS RUN Light C DF PW 229 . . . . . . . . .
JFS Operation PW 229 . . . . . . . . . . . . . . . .
Engine Controls andIndicators PW 229 . . . . . . . . . . . . . . . . . . . .
ENG CONT Switch PW 229 . . . . . . . . . . . .
AB RESET Switch C DF PW 229 . . . . . .
ENGINE FAULT Caution
Light PW 229 . . . . . . . . . . . . . . . . . . . . . .
Pilot Fault List Display(PFLD) PW 229 . . . . . . . . . . . . . . . . . . . .
SEC Caution Light PW 229 . . . . . . . . . . . .
EEC Caution Light PW 229 . . . . . . . . . . . .
BUC Caution Light PW 229 . . . . . . . . . . . .
MAX POWER Switch C DF PW 229 . . . .
RPM Indicator PW 229 . . . . . . . . . . . . . . . .
NOZ POS Indicator PW 229 . . . . . . . . . . . .
FTIT Indicator PW 229 . . . . . . . . . . . . . . . .
FUEL FLOW Indicator PW 229 . . . . . . . .
OIL Pressure Indicator PW 229 . . . . . . . .
HYD/OIL PRESS Warning
Light PW 229 . . . . . . . . . . . . . . . . . . . . . .
ENGINE Warning Light PW 229 . . . . . . .
Throttle PW 229 . . . . . . . . . . . . . . . . . . . . . .
Engine Operating
Characteristics PW 229 . . . . . . . . . . . . . . . .
Engine General PW 229 . . . . . . . . . . . . . . .
Ground Operations PW 229 . . . . . . . . . . . .
NonAB Operation in
Flight PW 229 . . . . . . . . . . . . . . . . . . . . . .
AB Operation in Flight PW 229 . . . . . . . .
SEC Operation PW 229 . . . . . . . . . . . . . . . .
Engine GE 129 . . . . . . . . . . . . . . . . . . . . . . . . . . . .
General Description GE 129 . . . . . . . . . . . . .
Engine Fuel/Control System GE 129 . . . . . .
Digital Electronic Control(DEC) GE 129 . . . . . . . . . . . . . . . . . . . . .
Afterburner Fuel Control(AFC) GE 129 . . . . . . . . . . . . . . . . . . . . . .
Main Engine Control (MEC) GE 129 . . .
Primary (PRI) Engine
Operation GE 129 . . . . . . . . . . . . . . . . . .

136
136
136
136
137
137
137
139
139
139
139
139
139
139
139
139
139
139
139
139
140
140
140
140
140
140
142
142
144
144
144
144
144
144
144

11

T.O. GR1F16CJ1
Hybrid (HYB) Engine
Operation GE 129 . . . . . . . . . . . . . . . . . .
Secondary Engine Control (SEC)
Operation GE 129 . . . . . . . . . . . . . . . . . .
Engine Fuel Boost Pump GE 129 . . . . . . .
Main Fuel Pump GE 129 . . . . . . . . . . . . . .
Afterburner (AB) Fuel
Pump GE 129 . . . . . . . . . . . . . . . . . . . . . .
Inlet Guide Vanes (IGV's) GE 129 . . . . . .
Variable Stator Vanes(VSV's) GE 129 . . . . . . . . . . . . . . . . . . . . .
Compressor Bleed Air GE 129 . . . . . . . . .
Exhaust Nozzle GE 129 . . . . . . . . . . . . . . . . . .
AB Flame Detector GE 129 . . . . . . . . . . . . . .
Engine Monitoring System(EMS) GE 129 . . . . . . . . . . . . . . . . . . . . . . . .
Engine Oil System GE 129 . . . . . . . . . . . . . . .
FUEL/OIL HOT Caution
Light GE 129 . . . . . . . . . . . . . . . . . . . . . .
Engine AntiIce System GE 129 . . . . . . . . . .
Engine ANTI ICE Switch GE 129 . . . . . .
INLET ICING Caution
Light GE 129 . . . . . . . . . . . . . . . . . . . . . .
Engine and Accessory Drive
Gearboxes GE 129 . . . . . . . . . . . . . . . . . . . .
Engine Alternator GE 129 . . . . . . . . . . . . . . .
Engine Ignition System GE 129 . . . . . . . . . .
Jet Fuel Starter (JFS) GE 129 . . . . . . . . . . . .
ENG & JET START Control
Panel GE 129 . . . . . . . . . . . . . . . . . . . . . . . .
JFS Switch C DF GE 129 . . . . . . . . . . . . .
JFS RUN Light C DF GE 129 . . . . . . . . .
JFS Operation GE 129 . . . . . . . . . . . . . . . .
Engine Controls andIndicators GE 129 . . . . . . . . . . . . . . . . . . . .
ENG CONT Switch GE 129 . . . . . . . . . . . .
SEC Caution Light GE 129 . . . . . . . . . . . .
EEC Caution Light GE 129 . . . . . . . . . . . .
BUC Caution Light GE 129 . . . . . . . . . . . .
MAX POWER Switch C DF GE 129 . . . .
AB RESET Switch C DF GE 129 . . . . . .
ENGINE FAULT Caution
Light GE 129 . . . . . . . . . . . . . . . . . . . . . .
Pilot Fault List Display(PFLD) GE 129 . . . . . . . . . . . . . . . . . . . .
RPM Indicator GE 129 . . . . . . . . . . . . . . .
NOZ POS Indicator GE 129 . . . . . . . . . . . .
FTIT Indicator GE 129 . . . . . . . . . . . . . . . .
FUEL FLOW Indicator GE 129 . . . . . . . .
OIL Pressure Indicator GE 129 . . . . . . . .

12

146
147
147
147
147
147
147
147
147
148
148
148
148
148
149
149
149
149
149
150
150
150
151
151
151
151
153
153
153
153
153
153
153
153
153
153
153
153

HYD/OIL PRESS Warning
Light GE 129 . . . . . . . . . . . . . . . . . . . . . .
ENGINE Warning Light GE 129 . . . . . . .
Throttle GE 129 . . . . . . . . . . . . . . . . . . . . . .
Engine Operating
Characteristics GE 129 . . . . . . . . . . . . . . .
Ground Operations GE 129 . . . . . . . . . . . .
NonAB Operation in Flight GE 129 . . . .
AB Operation GE 129 . . . . . . . . . . . . . . . . .
Inlet Thump GE 129 . . . . . . . . . . . . . . . . . .
SEC Operation GE 129 . . . . . . . . . . . . . . . .
Fire and Overheat Detection System . . . . . . . .
FIRE & OHEAT DETECT Test
Button C DF . . . . . . . . . . . . . . . . . . . . . . . .
Fuel System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Fuel Tank System . . . . . . . . . . . . . . . . . . . . . . .
Fuel Transfer System . . . . . . . . . . . . . . . . . . . .
Fuel Tank Vent and Pressurization
System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Engine Fuel Supply System . . . . . . . . . . . . . .
Fuel Quantity Indicating System . . . . . . . . .
Reservoir Fuel Level Sensing System . . . . .
Fuel Low Caution Lights . . . . . . . . . . . . . .
HUD FUEL Low/Bingo Indication . . . . . . . .
HUD TRP FUEL Warning . . . . . . . . . . . . . . .
FUEL/OIL HOT Caution Light . . . . . . . . . .
Fuel Tank Explosion Suppression
System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Refueling System . . . . . . . . . . . . . . . . . . . . . . .
Ground Refueling . . . . . . . . . . . . . . . . . . . . .
Air Refueling (AR) System . . . . . . . . . . . . .
NWS A/R DISC MSL STEP
Button . . . . . . . . . . . . . . . . . . . . . . . . . .
Air Refueling (AR) Status
Indicator . . . . . . . . . . . . . . . . . . . . . . . .
Environmental Control System (ECS) . . . . . . .
Electrical Failures . . . . . . . . . . . . . . . . . . . . . .
AirConditioning . . . . . . . . . . . . . . . . . . . . . . . .
Pressurization . . . . . . . . . . . . . . . . . . . . . . . . . .
AIR SOURCE Knob C DF . . . . . . . . . . . . .
TEMP Knob C DF . . . . . . . . . . . . . . . . . . . .
DEFOG Lever C DF . . . . . . . . . . . . . . . . . .
EQUIP HOT Caution Light . . . . . . . . . . . .
Cockpit Pressure Altimeter C DF . . . . . .
CABIN PRESS Caution Light . . . . . . . . . .
AntiG System . . . . . . . . . . . . . . . . . . . . . . . . . .
Electrical System . . . . . . . . . . . . . . . . . . . . . . . . . .
Main AC Power System . . . . . . . . . . . . . . . . . .
Overcurrent Sensing Contactors . . . . . . .

153
154
154
154
154
154
156
156
156
158
158
158
158
158
161
161
174
174
174
174
174
177
177
177
177
177
178
178
178
178
180
180
181
181
181
182
182
182
182
182
182
183

T.O. GR1F16CJ1Standby AC Power System . . . . . . . . . . . . . . . 183
Emergency AC Power System . . . . . . . . . . . . 183
DC Power System . . . . . . . . . . . . . . . . . . . . . . . 183
FLCS Power Supply . . . . . . . . . . . . . . . . . . . . . 183
External Power Provisions . . . . . . . . . . . . . . . 186
Electrical System Controls andIndicators . . . . . . . . . . . . . . . . . . . . . . . . . . . . 186
Electrical Power Distribution . . . . . . . . . . . . 186
Electrical System Normal Operation . . . . . . 186
Hydraulic System . . . . . . . . . . . . . . . . . . . . . . . . . 198
HYD PRESS Indicators and Warning
Light . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 198
HYD PRESS Indicators . . . . . . . . . . . . . . . 198
HYD/OIL PRESS Warning Light . . . . . . . 198
Emergency Power Unit (EPU) . . . . . . . . . . . . . . 198
EPU Controls and Indicators . . . . . . . . . . . . . 199
EPU Ground Safety Switch . . . . . . . . . . . . 199
EPU Switch . . . . . . . . . . . . . . . . . . . . . . . . . . 199
EPU Run Light . . . . . . . . . . . . . . . . . . . . . . . 199
HYDRAZN Light . . . . . . . . . . . . . . . . . . . . . 199
AIR Light . . . . . . . . . . . . . . . . . . . . . . . . . . . . 199
EPU/GEN Test Switch . . . . . . . . . . . . . . . . 199
EPU FUEL Quantity Indicator
C DF . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 199
Hydrazine Leak Detector . . . . . . . . . . . . . . 199
EPU Fired Indicator . . . . . . . . . . . . . . . . . 1102
EPU Operation . . . . . . . . . . . . . . . . . . . . . . . . 1102
Landing Gear (LG) System . . . . . . . . . . . . . . . . 1104
Main Landing Gear (MLG) . . . . . . . . . . . . . . 1104
Nose Landing Gear (NLG) . . . . . . . . . . . . . . 1104
Landing Gear Controls and Indicators . . . 1104
Landing Gear Handle . . . . . . . . . . . . . . . . 1104
Landing Gear Handle Down Permission
Button . . . . . . . . . . . . . . . . . . . . . . . . . . . 1104
DN LOCK REL Button . . . . . . . . . . . . . . . 1104
ALT GEAR Handle . . . . . . . . . . . . . . . . . . 1106
LG Warning Horn . . . . . . . . . . . . . . . . . . . 1106
HORN SILENCER Button . . . . . . . . . . . . 1106
TO/LDG CONFIG Warning Light . . . . . 1106
WHEELS Down Lights . . . . . . . . . . . . . . . 1106
Landing Gear WeightonWheels (WOW)
Switches . . . . . . . . . . . . . . . . . . . . . . . . . . 1106
Landing Gear Operation . . . . . . . . . . . . . . . . 1106
Nosewheel Steering (NWS) System . . . . . . . . . 1110
NWS Controls and Indicators . . . . . . . . . . . 1110
NWS A/R DISC MSL STEP Button . . . . 1110
NWS Light . . . . . . . . . . . . . . . . . . . . . . . . . . 1110
NWS FAIL Caution Light . . . . . . . . . . . . 1110
Nosewheel Steering (NWS) System D . . . . 1110

Wheel Brake System . . . . . . . . . . . . . . . . . . . . . .
Toe Brake System . . . . . . . . . . . . . . . . . . . . . .
Spin Down Braking System . . . . . . . . . . . . .
Brakes Channel Switch C DF . . . . . . . . . . .
Parking Brake C DF . . . . . . . . . . . . . . . . . . .
Antiskid System . . . . . . . . . . . . . . . . . . . . . . .
ANTISKID Switch C DF . . . . . . . . . . . .
ANTI SKID Caution Light . . . . . . . . . . . .
Speedbrake System . . . . . . . . . . . . . . . . . . . . . . .
SPD BRK Switch . . . . . . . . . . . . . . . . . . . .
SPEED BRAKE Position Indicator . . . .
Drag Chute System . . . . . . . . . . . . . . . . . . . . . . .
DRAG CHUTE Switch . . . . . . . . . . . . . . .
Arrestment System . . . . . . . . . . . . . . . . . . . . . . .
HOOK Switch . . . . . . . . . . . . . . . . . . . . . . .
HOOK Caution Light . . . . . . . . . . . . . . . .
Wing Flap System . . . . . . . . . . . . . . . . . . . . . . . .
Leading Edge Flaps (LEF's) . . . . . . . . . . . . .
LE FLAPS Switch . . . . . . . . . . . . . . . . . . .
FLCS LEF LOCK PFL . . . . . . . . . . . . . . .
Trailing Edge Flaps (TEF's)(Flaperons) . . . . . . . . . . . . . . . . . . . . . . . . . .
ALT FLAPS Switch . . . . . . . . . . . . . . . . . .
Flight Control System (FLCS) . . . . . . . . . . . . .
Digital Backup (DBU) . . . . . . . . . . . . . . . .
FLCS Limiters . . . . . . . . . . . . . . . . . . . . . . . . .
AOA/G Limiter . . . . . . . . . . . . . . . . . . . . . .
Roll Rate Limiter . . . . . . . . . . . . . . . . . . . .
Rudder Authority Limiter . . . . . . . . . . . .
Yaw Rate Limiter . . . . . . . . . . . . . . . . . . . .
FLCS Gains . . . . . . . . . . . . . . . . . . . . . . . . . . .
Cruise Gains . . . . . . . . . . . . . . . . . . . . . . . .
Takeoff and Landing Gains . . . . . . . . . . .
Standby Gains . . . . . . . . . . . . . . . . . . . . . .
FLCS Data Recorder C DF . . . . . . . . . . . . .
AngleofSideslip (AOS) Feedback
Function . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Gun Compensation . . . . . . . . . . . . . . . . . . . . .
Flight Control System (FLCS)
Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Stick . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Rudder Pedals . . . . . . . . . . . . . . . . . . . . . . .
MANUAL TRIM Panel . . . . . . . . . . . . . . .
MANUAL PITCH Override (MPO)
Switch . . . . . . . . . . . . . . . . . . . . . . . . . . .
STORES CONFIG Switch C DF . . . . . .
Low Speed Warning Tone . . . . . . . . . . . . .
FLT CONTROL Panel (FLCP) C DF . .

1110
1110
1111
1111
1111
1116
1117
1117
1117
1117
1117
1117
1118
1118
1118
1118
1118
1118
1118
1118
1120
1120
1120
1120
1121
1121
1121
1121
1121
1121
1127
1127
1127
1127
1127
1127
1127
1127
1130
1130
1130
1130
1130
1130

13

T.O. GR1F16CJ1FLCS Warning, Caution, and Indicator
Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FLCS Warning Light . . . . . . . . . . . . . . . . .
DBU ON Warning Light . . . . . . . . . . . . . .
TO/LDG CONFIG Warning Light . . . . .
FLCS FAULT Caution Light . . . . . . . . . .
BuiltIn Test (BIT) . . . . . . . . . . . . . . . . . . .
Pilot Fault List Display (PFLD) . . . . . . .
Trim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Autopilot . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Autopilot Operation . . . . . . . . . . . . . . . . . . . .
ADV MODE Switch . . . . . . . . . . . . . . . . . .
Stick Steering . . . . . . . . . . . . . . . . . . . . . . . . .
AOA System . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AOA Indicator . . . . . . . . . . . . . . . . . . . . . . .
AOA Indexer . . . . . . . . . . . . . . . . . . . . . . . .
HUD AOA Display . . . . . . . . . . . . . . . . . . .
Air Data System . . . . . . . . . . . . . . . . . . . . . . . . . .
Air Data Probes . . . . . . . . . . . . . . . . . . . . .
AOA Transmitters . . . . . . . . . . . . . . . . . . .
Total Temperature Probe . . . . . . . . . . . . .
Static Pressure Ports . . . . . . . . . . . . . . . . .
Probe Heat Monitor . . . . . . . . . . . . . . . . . .
PROBE HEAT Switch . . . . . . . . . . . . . . . .
Pneumatic Sensor Assembly (PSA) . . . . . .
Central Air Data Computer (CADC) . . . . .
CADC Caution Light . . . . . . . . . . . . . . . . .
Warning, Caution, and Indicator Lights . . . .
Voice Message System (VMS) . . . . . . . . . . .
VOICE MESSAGE Switch C DF . . . . .
MASTER CAUTION Light . . . . . . . . . . . . . .
Pilot Fault List Display (PFLD) . . . . . . . . .
Caution Light Panel . . . . . . . . . . . . . . . . . . . .
MAL & IND LTS Test Button . . . . . . . . .
Lighting System . . . . . . . . . . . . . . . . . . . . . . . . . .
Exterior Lighting . . . . . . . . . . . . . . . . . . . . . .
Anticollision Strobe Light . . . . . . . . . . . .
Position/Formation Lights . . . . . . . . . . . .
Air Refueling Lights . . . . . . . . . . . . . . . . .
Vertical TailMounted Floodlight . . . . . .
Landing and Taxi Lights . . . . . . . . . . . . .
Interior Lighting . . . . . . . . . . . . . . . . . . . . . . .
PRIMARY CONSOLES Knob . . . . . . . . .
PRIMARY INST PNL Knob . . . . . . . . . . .
DATA ENTRY DISPLAY Knob . . . . . . . .
MAL & IND LTS Switch . . . . . . . . . . . . . .
FLOOD INST PNL Knob . . . . . . . . . . . . .
FLOOD CONSOLES Knob . . . . . . . . . . .

14

1130
1130
1132
1132
1132
1135
1135
1135
1135
1136
1136
1136
1137
1137
1138
1138
1138
1138
1138
1138
1138
1138
1140
1140
1140
1140
1140
1145
1146
1146
1146
1147
1147
1148
1148
1148
1148
1148
1154
1154
1154
1155
1155
1155
1155
1155
1155

NVIS Switch PX III . . . . . . . . . . . . . . . . . .
HandsOn Blackout (HOBO)
Switch PX III . . . . . . . . . . . . . . . . . . . . . .
Utility Light . . . . . . . . . . . . . . . . . . . . . . . .
Cockpit Spotlights . . . . . . . . . . . . . . . . . . .
Escape System . . . . . . . . . . . . . . . . . . . . . . . . . . .
Canopy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Canopy Controls and Indicators . . . . . . . . .
MANUAL CANOPY CONTROL
Handcrank . . . . . . . . . . . . . . . . . . . . . . .
Canopy Handle C DF . . . . . . . . . . . . . . . .
Canopy Jettison . . . . . . . . . . . . . . . . . . . . . . . .
Ejection Seat . . . . . . . . . . . . . . . . . . . . . . . . . .
Ejection Handle . . . . . . . . . . . . . . . . . . . . .
SHOULDER HARNESS Knob . . . . . . . .
EMERGENCY MANUAL CHUTE
Handle . . . . . . . . . . . . . . . . . . . . . . . . . . .
Ejection Safety Lever . . . . . . . . . . . . . . . .
SEAT NOT ARMED Caution Light . . . .
SEAT ADJ Switch . . . . . . . . . . . . . . . . . . .
Shoulder Harness Straps/Parachute
Risers . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Inertia Reel Straps . . . . . . . . . . . . . . . . . .
Lapbelt . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Survival Kit . . . . . . . . . . . . . . . . . . . . . . . . .
RADIO BEACON Switch . . . . . . . . . . . . .
Emergency Oxygen . . . . . . . . . . . . . . . . . .
Ejection Seat Operation . . . . . . . . . . . . . . . .
Ejection Mode Envelopes . . . . . . . . . . . . . . .
Ejection Sequence Times . . . . . . . . . . . . . . . .
Canopy Jettison/Seat Ejection . . . . . . . . . . .
Ejection Seat Performance . . . . . . . . . . . . . .
Oxygen System PX III . . . . . . . . . . . . . . . . . . . . .
Onboard Oxygen Generating
System (OBOGS) . . . . . . . . . . . . . . . . . . . .
OBOGS Backup Oxygen
Supply (BOS) . . . . . . . . . . . . . . . . . . . . . . . .
OBOGS Monitor . . . . . . . . . . . . . . . . . . . . . . .
Emergency Oxygen Supply (EOS) . . . . . . .
OBOGS BuiltInTests (BIT's) . . . . . . . . . . .
OBOGS PowerUp BIT . . . . . . . . . . . . . . . . .
OBOGS Periodic BIT . . . . . . . . . . . . . . . . . . .
OBOGS Manual BIT . . . . . . . . . . . . . . . . . . .
OBOGS Pneumatic BIT . . . . . . . . . . . . . . . . .
Pressure Breathing for G (PBG) . . . . . . . . .
Oxygen System Controls andIndicators . . . . . . . . . . . . . . . . . . . . . . . . . . .
Oxygen System Schematic . . . . . . . . . . . . . .

1155
1157
1157
1157
1157
1157
1157
1157
1157
1161
1161
1161
1161
1161
1161
1161
1161
1161
1161
1161
1164
1164
1164
1164
1164
1164
1164
1164
1173
1173
1173
1173
1173
1173
1173
1173
1173
1174
1174
1174
1174

T.O. GR1F16CJ1Oxygen System PX II . . . . . . . . . . . . . . . . . . . . .
Liquid Oxygen System . . . . . . . . . . . . . . . . . .
Pressure Breathing for G (PBG) . . . . . .
Oxygen System Controls andIndicators . . . . . . . . . . . . . . . . . . . . . . . .
Oxygen Duration . . . . . . . . . . . . . . . . . . . .
Emergency Oxygen System . . . . . . . . . . . . .
Oxygen System Schematic . . . . . . . . . . . . . .
Communications, Navigation, and IFF(CNI) Equipment . . . . . . . . . . . . . . . . . . . . . .
Upfront Controls . . . . . . . . . . . . . . . . . . . . . . .
DED/CNI Readouts . . . . . . . . . . . . . . . . . . . .
CNI Readout/DED Page Summary . . . .
IFF Control Panel C DF PX III . . . . . . . . . .
C & I (Communications and IFF)
Knob . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
IFF MASTER Knob . . . . . . . . . . . . . . . . . .
IFF M4 CODE Switch . . . . . . . . . . . . . . .
IFF MODE 4 REPLY Switch . . . . . . . . . .
IFF MODE 4 MONITOR Switch . . . . . .
IFF ENABLE Switch . . . . . . . . . . . . . . . . .
IFF MODE 1/MODE 3 Selector
Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AUX COMM Panel C DF PX II . . . . . . . . . .
CNI (Communications, Navigation,and IFF) Knob . . . . . . . . . . . . . . . . . . . .
IFF MASTER Knob . . . . . . . . . . . . . . . . . .
M4 CODE Switch . . . . . . . . . . . . . . . . . . .
IFF MODE 4 REPLY Switch . . . . . . . . . .
IFF MODE 4 MONITOR Switch . . . . . .
TACAN Function Switch . . . . . . . . . . . . .
TACAN CHANNEL Selector Levers . . .
AUDIO 1 Control Panel . . . . . . . . . . . . . . . . .
COMM 1 (UHF) Power Knob . . . . . . . . .
COMM 1 (UHF) Mode Knob . . . . . . . . . .
COMM 2 (VHF) Power Knob . . . . . . . . . .
COMM 2 (VHF) Mode Knob . . . . . . . . . .
SECURE VOICE Knob . . . . . . . . . . . . . . .
MSL Tone Knob . . . . . . . . . . . . . . . . . . . . .
TF Tone Knob . . . . . . . . . . . . . . . . . . . . . . .
THREAT Tone Knob . . . . . . . . . . . . . . . . .
AUDIO 2 Control Panel . . . . . . . . . . . . . . . . .
ILS Power Knob . . . . . . . . . . . . . . . . . . . . .
TACAN Power Knob . . . . . . . . . . . . . . . . .
INTERCOM Knob . . . . . . . . . . . . . . . . . . .
HOT MIC CIPHER Switch . . . . . . . . . . .
ANT SEL Panel C DF . . . . . . . . . . . . . . . . . .
IFF ANT SEL Switch C DF . . . . . . . . . .
UHF ANT SEL Switch C DF . . . . . . . . .
Communications System . . . . . . . . . . . . . . . . . .

1177
1177
1177
1177
1177
1177
1177
1182
1182
1182
1185
1185
1186
1186
1186
1186
1186
1187
1187
1187
1187
1187
1187
1188
1188
1188
1188
1188
1188
1188
1189
1189
1189
1189
1189
1189
1190
1190
1190
1190
1190
1190
1190
1191
1191

Antenna Locations . . . . . . . . . . . . . . . . . . . . .
Have Quick (HQ) System . . . . . . . . . . . . . . .
Single Word of Day (WOD) . . . . . . . . . . .
Single WOD Transfer . . . . . . . . . . . . . . . .
Multiple Word of Day (MWOD) . . . . . . .
Automatic MWOD Loading PX III . . . . .
TimeofDay (TOD) Transmission . . . . .
TimeofDay (TOD) Reception PX III . . .
TimeofDay (TOD) Reception PX II . . .
Net Number . . . . . . . . . . . . . . . . . . . . . . . . .
Antijamming (AJ) Mode Operation . . . .
Have Quick (HQ) UHF Radio . . . . . . . . . . . .
Operation of the UHF Radio on
Upfront Controls . . . . . . . . . . . . . . . . . .
UHF Radio Backup Control Panel . . . . .
Function Knob . . . . . . . . . . . . . . . . . . . . . .
Mode Knob PX II . . . . . . . . . . . . . . . . . . . .
Mode Knob PX III . . . . . . . . . . . . . . . . . . . .
CHAN Knob . . . . . . . . . . . . . . . . . . . . . . . . .
Manual Frequency Knobs . . . . . . . . . . . .
A32T Knob PX II . . . . . . . . . . . . . . . . . .
A32 Knob PX III . . . . . . . . . . . . . . . . . . . .
VOL Knob . . . . . . . . . . . . . . . . . . . . . . . . . .
SQUELCH Switch . . . . . . . . . . . . . . . . . . .
TONE Button PX II . . . . . . . . . . . . . . . . . .
CHAN Display PX III . . . . . . . . . . . . . . . .
Frequency/Status Display PX III . . . . . .
STATUS Button PX III . . . . . . . . . . . . . . .
TTONE Switch PX III . . . . . . . . . . . . . . . .
TEST DISPLAY Button PX III . . . . . . . .
ZERO Switch PX III . . . . . . . . . . . . . . . . . .
FILL Connector PX III . . . . . . . . . . . . . . .
VHF Radio . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Operation of the VHF Radio on
Upfront Controls . . . . . . . . . . . . . . . . . .
Zeroize Switch . . . . . . . . . . . . . . . . . . . . . . . . . . .
Navigation Systems . . . . . . . . . . . . . . . . . . . . . .
Embedded Global Positioning and Inertial
Navigation Set (EGI) PX III . . . . . . . . . . .
Global Positioning System (GPS) PX II . . .
Inertial Navigation Set (INS) PX II . . . . . .
Tactical Air Navigation (TACAN)
System . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Operation of the TACAN on Upfront
Controls . . . . . . . . . . . . . . . . . . . . . . . . . .
Terrain Following Radar (TFR) System . . . . .
AIFF System . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AIFF Transponder PX II . . . . . . . . . . . . . . . .
AIFF Transponder PX III . . . . . . . . . . . . . . .
Operation of the Transponder on
Upfront Controls PX II . . . . . . . . . . . . .
Operation of the Transponder on
Upfront Controls PX III . . . . . . . . . . . .

1191
1191
1191
1191
1191
1193
1193
1193
1193
1194
1194
1194
1194
1196
1196
1196
1196
1197
1197
1197
1197
1197
1197
1197
1197
1197
1198
1198
1198
1198
1198
1198
1198
1199
1199
1199
1200
1200
1200
1200
1200
1200
1201
1201
1201
1202

15

T.O. GR1F16CJ1AIFF Interrogator . . . . . . . . . . . . . . . . . . . . . .
Operation of the Interrogator on
Upfront Controls PX II . . . . . . . . . . . . .
Operation of the Interrogator on
Upfront Controls PX III . . . . . . . . . . . .
AIFF Controls and Indicators . . . . . . . . . . .
IFF IDENT Button . . . . . . . . . . . . . . . . . .
IFF (Mode 4) Caution Light . . . . . . . . . . .
UHF VHF IFF Transmit
Switch PX II . . . . . . . . . . . . . . . . . . . . . .
Target Management Switch PX III . . . . .
Instrument Landing System (ILS) . . . . . . . . .
ILS Controls . . . . . . . . . . . . . . . . . . . . . . . .
Operation of the ILS on Upfront
Controls . . . . . . . . . . . . . . . . . . . . . . . . . .
MRK BCN Light . . . . . . . . . . . . . . . . . . . . .
INSTR MODE Select Panel . . . . . . . . . . . . . . . .
INSTR HDG Knob C DF . . . . . . . . . . . . .
INSTR MODE Knob . . . . . . . . . . . . . . . . .
Navigation Aids and Display . . . . . . . . . .
Instrument Modes . . . . . . . . . . . . . . . . . . .
Flight Instruments . . . . . . . . . . . . . . . . . . . . . . .
Altimeter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Airspeed/Mach Indicator . . . . . . . . . . . . . . . .
Standby Attitude Indicator (SAI) . . . . . . . .
Vertical Velocity Indicator (VVI) . . . . . . . . .
Magnetic Compass . . . . . . . . . . . . . . . . . . . . .
Attitude Director Indicator (ADI) . . . . . . . .
Horizontal Situation Indicator (HSI) . . . . .
Clock . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Crash Survivable Flight Data
Recorder (CSFDR) . . . . . . . . . . . . . . . . . . . . .
F16D Aircraft . . . . . . . . . . . . . . . . . . . . . . . . . . .
Engine Controls DR . . . . . . . . . . . . . . . . . . . .
Throttle DR . . . . . . . . . . . . . . . . . . . . . . . . .
ENG CONT Switch DR . . . . . . . . . . . . . . .
Fuel System . . . . . . . . . . . . . . . . . . . . . . . . . . .
Fuel Control Panel DR . . . . . . . . . . . . . . .
Speedbrake System . . . . . . . . . . . . . . . . . . . .
SPD BRK Switch DR . . . . . . . . . . . . . . . . .
Drag Chute System . . . . . . . . . . . . . . . . . . . .
DRAG CHUTE Switch DR . . . . . . . . . . . .
Landing Gear (LG) System . . . . . . . . . . . . . .
DN LOCK REL Button DR . . . . . . . . . . .
Flight Control System . . . . . . . . . . . . . . . . . .
Stick Control Operation . . . . . . . . . . . . . . . .
STICK CONTROL Switch DF . . . . . . . . .
Stick Indicator DR . . . . . . . . . . . . . . . . . . .
Stick OVRD Lights D . . . . . . . . . . . . . . . .

16

1203
1204
1204
1204
1204
1204
1204
1205
1205
1205
1206
1206
1206
1206
1206
1206
1206
1215
1215
1216
1216
1216
1216
1216
1216
1217
1217
1219
1219
1219
1219
1219
1219
1219
1219
1219
1219
1219
1219
1219
1219
1220
1220
1220

Escape System D . . . . . . . . . . . . . . . . . . . . . .
EJECTION MODE SEL Handle DR . . .
Communications System . . . . . . . . . . . . . . .
INSTR Mode Select Panel D . . . . . . . . . . . .
Flight Instruments DR . . . . . . . . . . . . . . . . .
Accelerometer DR . . . . . . . . . . . . . . . . . . . .
Aft Station Interface Unit
(ASIU) DR PX III . . . . . . . . . . . . . . . . . . . .
STICK CONTROL Switch DR PX III . . .
ASIU Switch DR PX III . . . . . . . . . . . . . . .
Multifunction Display (MFD) DR . . . . . . . .
Throttle and Stick Switches/Controls
DR . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ARMT CONSENT Switch DR . . . . . . . . . . .
Aft Seat HUD Monitor (ASHM) DR . . . . . .
Servicing Diagram . . . . . . . . . . . . . . . . . . . . .

1220
1220
1221
1221
1221
1221
1221
1221
1222
1222
1222
1222
1222
1222

THE AIRCRAFTThe F16C is a singleengine, singleseat, multirole
tactical fighter with full airtoair and airtosurface
combat capabilities. The F16D is a twoseat(tandem) version and performs the secondary role ofa trainer. The fuselage is characterized by a large
bubble canopy, forebody strakes, and an underfuselage engine air inlet. The wing and tail surfaces arethin and feature moderate aft sweep. The wing has
automatic leading edge flaps which enhance
performance over a wide speed range. Flaperons aremounted on the trailing edge of the wing andcombine the functions of flaps and ailerons. The
horizontal tails have a small negative dihedral andprovide pitch and roll control through symmetrical/
differential deflection. The vertical tail, augmented
by twin ventral fins, provides directional stability.
All flight control surfaces are actuated hydraulicallyby two independent hydraulic systems and are
directed by signals through a flybywire system.
The fire control system includes a fire control radarwith search and tracking capability; two multifunc
tion displays (MFD's) for presentation of navigation,
radar, weapons, and other information; and aheadup display (HUD). Upfront controls provide a
central control point for fire control, communica
tions, navigation, and IFF functions. A stores
management system (SMS) presents a visualdisplay on a selected MFD for inventory, control, andrelease of all stores. Basic armament includes afuselagemounted multibarrel 20 mm gun and an
airtoair missile on each wingtip. Additional storesof various types can be carried on pylons mounted
under the wings and on the fuselage.

T.O. GR1F16CJ1AIRCRAFT GENERAL DATAAIRCRAFT GROSS WEIGHT (A/C GW)

Refer to figure 11. The aircraft approximate
dimensions are:

The GW of the aircraft including pilot, oil, two tip
AIM120 missiles, and a full load of 20 mm
ammunition is approximately 21,000 pounds andwith full internal JP8 fuel 28,200 pounds.

D Span-including missilefins

D Length-including noseprobe

= 32 feet 10 inches= 49 feet 5.2 inches

D Height-top of vertical tail= 16 feet 10 inches

D Height-top of canopy= 9 feet 4 inches

D Tread= 7 feet 9 inches

D Wheelbase= 13 feet 2 inches

Refer to TURNING RADIUS AND GROUND
CLEARANCE, Section II.

C

The GW of the aircraft including pilot, oil,
two tip AIM120 missiles, and a full load of 20 mm
ammunition is approximately 21,600 pounds andwith full internal JP8 fuel 27,500 pounds.
PX II DPX III D The GW of the aircraft including pilot, oil,
two tip AIM120 missiles, and a full load of 20 mm
ammunition is approximately 21,900 pounds andwith full internal JP8 fuel 27,800 pounds.
PX III CFT's add approximately 900 pounds whenempty and hold approximately 3000 pounds of JP8
fuel.

These GW's are approximate and shall not be used forcomputing aircraft performance. For maximum GW
limitations, refer to Section V, OPERATING LIMITATIONS. For detailed information, refer to T.O.
GR1F16CJ11, Part I.
COCKPIT ARRANGEMENTAIRCRAFT GENERAL ARRANGEMENT
Refer to figure 12 for general arrangement andoverall view of the aircraft.

Refer to figure 13. The cockpit is conventional except
for the seat, which is reclined 30 degrees, and the
stick, which is mounted on the right console. The
cockpit contains no circuit breakers.

17

T.O. GR1F16CJ1General Data
AIM-9 MISSILEMISSILE LAUNCHER
15.16

13.32
OUTBOARD STORES LOCATIONSFLAPERON
40

370-GALLON FUEL TANK

37.39
69.63

28.63

95.82
30.63
195.52

18 FT 4.2 IN.
SPEEDBRAKE HINGE
29.72

70.87

8.45

135.84
190.43

12.05

44.42

SHADED AREA REPRESENTS PX III

D AIRCRAFT

49 FT 5.2 IN.
OVERALL LENGTH

46.80

82.05

15.35
29.08

108.621595.0 101.0

19.5
16 FT 10.0 IN.
OVERALL
HEIGHT

22.20
STATIC GROUND LINE
49.81
NOTE: Dimensions are in inches unless specified otherwise.

GR1F-16CJ-1-0004-1A37

Figure 11.(Sheet 1)18

T.O. GR1F16CJ1

WINGS
Area
Span
Aspect Ratio
Taper Ratio
Sweep (LE)DihedralAirfoil
Incidence
Twist
At BL 54.0
At BL 180.0
Flaperon Area
LEF Area
HORIZONTAL TAILS
Area
Aspect Ratio
Taper Ratio
Sweep (LE)DihedralAirfoilAt Root
At TipVERTICAL TAIL
Area
Aspect Ratio
Taper Ratio
Sweep (LE)
AirfoilAt Root
At TipRudder Area

300 Sq Ft
30 Ft
3.0
0.2275400NACA 64A204
00
3
31.32 Sq Ft
36.71 Sq Ft

VENTRAL FIN (EACH)
Area
Span
Aspect Ratio
Taper Ratio
Sweep (LE)Dihedral (Cant)
AirfoilAt Root
At Tip8.03 Sq Ft
23.356 In. Theo (27.5 In. Actual)
0.472 (Theo)
0.760 (Theo)30
15 Outboard
3.886% Modified Wedge
Constant 0.03R

SPEEDBRAKES
Area (4 Element Clamshell)

63.70 Sq Ft
2.114
0.390 (Theo)40-106% Biconvex
3.5% Biconvex
54.75 Sq Ft
1.294
0.437
47.5
5.3% Biconvex
3.0% Biconvex
11.65 Sq Ft

LANDING GEAR (LG)Main Gear (MLG)
Tire Size
StrokeStatic Rolling Radius
Nose Gear (NLG)
Tire Size
StrokeStatic Rolling Radius

14.26 Sq Ft (3.565 Sq Ft Ea)

27.75 x 8.75-14.5 24 Ply
10.5 In.
11.9 In.
17.9 x 6.7R8 22 Ply
10.0 In.
7.5 In.

ENGINE
F100PW229
ThrustEngine Compressor
Face Diameter
Engine Length

29,000 Lb Class
34.8 In.
208 In.

F110-GE-129
ThrustEngine Compressor
Face Diameter
Engine Length

29,500 Lb Class
35.8 In.
183.76 In.

32 FT 10 IN. *OVERALL SPAN W/MISSILES
31 FTOVERALL SPAN W/O MISSILES

10
46.5
93.0
* Add 3 inches for AIM-120 missiles
GR1F-16CJ-1-0004-2A37

Figure 11.(Sheet 2)

19

8

11

11

8

25

15

17
27

23

13

C

For antenna locations, refer
to ANTENNA LOCATIONS,
this section.

26
24

D

8
5

9 10

11

T.O. GR1F16CJ1

110

General Arrangement (Typical)
16
14 15

22
17
18

19
20

30

21

28
29
31

12
38

6 7

39

37

32
33

34 17
36

40

41

34

42

35

17

2

1

4
1. Air Data (Pitot) Probe
2. Nose Radome
3. AOA Probe
4. AOA Transmitter
5. Forward Electronic Equipment Bay
6. Cockpit Pressure Regulator
7. Cockpit Pressure Safety Valve
8. HUD/ASHM
9. Instrument Panel
10. Canopy (Movable)
11. Ejection Seat
12. Canopy (Fixed)
13. AIM-9 Missile
14. Missile Launcher
15. AR and Formation Light
16. Position/Formation Light

3

63

62 61
60
59

54
55
56
58 57

17. Static Discharger
18. LEF Drive Unit
19. Hydraulic Reservoir
20. AR Slipway
21. FLCS Accumulators
22. Anticollision Strobe Light Power Supply
23. Vertical Tail Attachment Fittings
24. Vertical Tail
25. Vertical Tail-Mounted Floodlight
26. Anticollision Strobe Light
27. Rudder
28. Position Light
29. Drag Chute Fairing
30. Rudder ISA
31. Turbofan Engine
32. Speedbrake

Figure 12.

52
53

51

44

50 49 44
48

47

33. Speedbrake Actuator
34. Chaff/Flare Dispenser
35. Horizontal Tail ISA
36. Horizontal Tail
37. Formation Light
38. Flaperon ISA
39. Hook
40. Flaperon
41. Wing Structural Box
42. Fixed Trailing Edge Panel
43. Asymmetry Brake
44. LEF Rotary Actuator
45. LEF Torque Shaft
46. LEF
47. Wing Attachment Fittings
48. MLG

46

45 44
44 43

14

13
16

49. MLG Door
50. LEF Angle Drive Gearbox
51. Air-Conditioning Package
52. M61A1 20MM Gun
53. Ammunition Drum
54. EPU Nitrogen Bottle
55. Gun Port
56. NLG Door
57. Position Light
58. NLG
59. Engine Air Inlet
60. Left Console
61. Throttle
62. Strake
63. Lower Equipment Compartment
GR1F-16CJ-1-0005X37

T.O. GR1F16CJ1

Cockpit Arrangement C DF (Typical) PX IIDFDFINSTRUMENT
PANEL

LEFT
AUXILIARYCONSOLERIGHT
AUXILIARYCONSOLESTICK

THROTTLEDFLEFTCONSOLERIGHTCONSOLEDF

GR1F-16CJ-1-0006A37

Figure 13.(Sheet 1)

111

T.O. GR1F16CJ1

Cockpit Arrangement C DF (Typical) PX IILEFT CONSOLE

23
22
24

1
2

21

20

4

18 19

5

17

15 16

13

14

7

12

9

11

3

6

8

DF
10
1. UHF Backup Control Panel
2. AUDIO 1 Control Panel
3. AUDIO 2 Control Panel
4. ECM Control Panel
5. AVTR Control Panel
6. EXT LIGHTING Control Panel
7. MANUAL TRIM Panel
8. G-Suit Hose Connection
9. ANTI G TEST Button
10. DF STICK CONTROL Switch
11. Stowage
12. TEST Switch Panel
13. DEFOG Lever
14. FLT CONTROL Panel
15. Fuel Control Panel
16. AUX COMM Panel
17. CANOPY JETTISON T-Handle
18. EPU Control Panel
19. ELEC Control Panel
20. Throttle FRICTION Control
21. ENG & JET START Control Panel
22. MANUAL PITCH Override Switch
23. CHAFF/FLARE Dispense Button
24. Throttle

DF
GR1F-16CJ-1-0007A37

Figure 13.(Sheet 2)112

T.O. GR1F16CJ1

Cockpit Arrangement C DF (Typical) PX IILEFT AUXILIARY CONSOLERIGHT AUXILIARY CONSOLE

DF

18

19

1

2

3

4
5

17

1

6

16
15
12

14
13

7

10

4

3

8

11

2

6

5
7
8

9

1. EMER STORES JETTISON Button (Covered)
2. WHEELS Down Lights (Green)
3. HOOK Switch (Lever Lock)
4. ANTI-SKID Switch
5. DN LOCK REL Button
6. LG Handle Down Permission Button
7. LG Handle
8. LG Handle Warning Light (Red)
9. LANDING TAXI LIGHTS Switch
10. CMDS Control Panel
11. RF Transmit Panel
12. RWR Aux Control Panel
13. ALT GEAR Handle
14. ALT GEAR Reset Button
15. SPEED BRAKE Position Indicator
16. STORES CONFIG Switch
17. HORN SILENCER Button
18. GND JETT ENABLE Switch (Lever Lock)
19. BRAKES Channel Switch

910

1. Magnetic Compass
2. FUEL Quantity Indicator
3. Pilot Fault List Display
4. System A HYD PRESS Indicator
5. System B HYD PRESS Indicator
6. Caution Light Panel
7. LIQUID OXYGEN Quantity Indicator
8. EPU FUEL Quantity Indicator
9. Cockpit Pressure Altimeter
10. Clock

GR1F-16CJ-1-0008A37

Figure 13.(Sheet 3)

113

T.O. GR1F16CJ1

Cockpit Arrangement C DF (Typical) PX IIINSTRUMENT PANEL

DF

1. HUD Combiner GlassDF
2. AOA Indexer
3. AR Status/NWS Indicator422
4. Integrated Control Panel (ICP)
5. Standby Attitude Indicator
6. FUEL FLOW Indicator
41
7. Data Entry Display
40
8. ENG FIRE and ENGINE Warning
39
Lights (Red)38
36
9. HYD/OIL PRESS Warning Light (Red)37
10. FLCS and DBU ON Warning
35
34
Lights (Red)28
11. TO/LDG CONFIG Warning Light (Red)33
32
12. CANOPY Warning Light (Red)31
13. Right MFD
22
14. OIL Pressure Indicator29
15. NOZ POS Indicator30
16. RPM Indicator
2330. Autopilot ROLL Switch
17. FTIT Indicator
31. ADV MODE Switch
18. Vertical Velocity Indicator
32. MASTER ARM Switch
21
19. MRK BCN Light
33. C ALT REL Button
20. FUEL QTY SEL Panel
34. LASER ARM Switch
21. Rudder PEDAL ADJ Knob
35. IFF IDENT Button
22. AOA Indicator
36. DRAG CHUTE Switch
23. INSTR MODE Select Panel
37. RWR Prime Control Panel
24. Airspeed/Mach Indicator
38. Azimuth Display/Indicator
25. Attitude Director Indicator
39. F-ACK Button
26. Horizontal Situation Indicator
40. TF FAIL Warning Light (Red)
27. Altimeter
41. MASTER CAUTION Light (Amber)
28. Left MFD
42. ECM Light (Green), DF ECM (Green)
and OVRD (Amber) Light
29. Autopilot PITCH Switch

Figure 13.(Sheet 4)114

3

7
6 8
9

1
4

24

12
14
15
16

5
13

27
25

18
19

26 20

36
34

10
11

17

DF
39

32
31
30

29

GR1F-16CJ-1-0009A37

T.O. GR1F16CJ1

Cockpit Arrangement C DF (Typical) PX IIRIGHT CONSOLE

17
16
DF
18

19

1
2

4

3

5

6

15
14

10
13

12

7
9

8

11

DF

1. SNSR PWR Control Panel
2. HUD Control Panel
3. Interior LIGHTING Control Panel
4. NUCLEAR CONSENT Switch (Guarded)
5. PLAIN Cipher Switch
6. AIR COND Control Panel
7. ENGINE ANTI-ICE Switch
8. Utility Light
9. ANT SEL Panel
10. AVIONICS POWER Panel
11. OXYGEN REGULATOR Panel
12. DTU Control Panel
13. Stowage
14. Oxygen/Communication Hookup
15. VOICE MESSAGE Switch
16. ZEROIZE Switch
17. Stick
18. DF NWS Control Button/Indicator (Green)
19. SEAT ADJ Switch

GR1F-16CJ-1-0010A37

Figure 13.(Sheet 5)

115

T.O. GR1F16CJ1

Cockpit Arrangement DR (Typical) PX IIINSTRUMENT
PANEL

LEFT
AUXILIARYCONSOLERIGHT
AUXILIARYCONSOLESTICK

THROTTLELEFT
CONSOLERIGHTCONSOLE

GR1F-16CJ-1-0011A37

Figure 13.(Sheet 6)116

T.O. GR1F16CJ1

Cockpit Arrangement DR (Typical) PX IILEFT CONSOLE

7
9
8

1
2

6

10
5
11

4

3

1. AUDIO 1 Control Panel
2. AUDIO 2 Control Panel
3. G-Suit Hose Connection
4. ANTI G TEST Button
5. MAL & IND LTS Test Button
6. FUEL MASTER Switch
7. MANUAL PITCH Override Switch
8. ENG CONT Switch
9. Throttle
10. CANOPY JETTISON T-Handle
11. Stowage

GR1F-16CJ-1-0012A37

Figure 13.(Sheet 7)

117

T.O. GR1F16CJ1

Cockpit Arrangement DR (Typical) PX IILEFT AUXILIARY CONSOLE

1

15

2

3

RIGHT AUXILIARY CONSOLE

4
7

14

5

12
10

6
1

9

11
8

6

13

3
4

1. EMER STORES JETTISON Button (Covered)
2. WHEELS Down Lights (Green)
3. HOOK Switch (Lever Lock)
4. DN LOCK REL Button
5. LG Handle Down Permission Button
6. LG Handle Warning Light (Red)
7. LG Handle
8. ALT GEAR Handle
9. ALT GEAR Reset Button
10. Blank Panel
11. SPEED BRAKE Position Indicator
12. ALT FLAPS Switch (Lever Lock)
13. HORN SILENCER Button
14. GND JETT ENABLE Switch (Lever Lock)
15. DRAG CHUTE Switch

Figure 13.(Sheet 8)118

2

5

1. Clock
2. FUEL Quantity Indicator
3. System B HYD PRESS Indicator
4. System A HYD PRESS Indicator
5. EJECTION MODE SELECT Handle
6. Caution Light Panel
GR1F-16CJ-1-0013A37

T.O. GR1F16CJ1

Cockpit Arrangement DR (Typical) PX IIINSTRUMENT PANEL

1. MASTER CAUTION Light (Amber)
2. ENG FIRE and ENGINE Warning
Lights (Red)3. HYD/OIL PRESS Warning Light (Red)
4. FLCS and DBU ON Warning Lights (Red)
5. AR Status/NWS Indicator
6. TO/LDG CONFIG Warning Light (Red)
7. CANOPY Warning Light (Red)
8. Right MFD
9. VIDEO SEL Switch
10. Standby Attitude Indicator
11. FUEL FLOW Indicator
12. OIL Pressure Indicator
13. FAULT ACK Button
14. NOZ POS Indicator
15. RPM Indicator
16. FTIT Indicator
17. Pilot Fault List Display
18. Accelerometer
19. Altimeter
20. Vertical Velocity Indicator
21. Attitude Director Indicator
22. Horizontal Situation Indicator
23. MRK BCN Light
24. Rudder PEDAL ADJ Knob
25. INSTR MODE Select Panel
26. AOA Indicator
27. Airspeed/Mach Indicator
28. ARMT CONSENT Switch (Guarded)

39

32

33

38
37

40 41

35

5
8
11

36

7 9
12

10

28
26
25

21
22

13
14
15

27 19 1831

29. IFF IDENT Button
30. Data Entry Display
31. Integrated Keyboard Panel (IKP)
32. Stick Indicator
33. OVRD Light (Amber)
34. Blank
35. Azimuth Display/Indicator
36. Left MFD
37. AOA Indexer
38. TF FAIL Warning Light (Red)39. Handgrip
40. ECM Light (Green)
41. Aft Seat HUD Monitor

46

1

34

30 29

2 3

17
20

16

23

24

GR1F-16CJ-1-0014A37

Figure 13.(Sheet 9)

119

T.O. GR1F16CJ1

Cockpit Arrangement DR (Typical) PX IIRIGHT CONSOLE

7
8

6
1

5

2
4

3

1. Interior LIGHTING Control Panel
2. OXYGEN REGULATOR Panel
3. Utility Light
4. Oxygen/Communications Hookup
5. Stowage
6. NWS Control Button/Indicator (Green)
7. Stick
8. SEAT ADJ Switch

GR1F-16CJ-1-0015A37

Figure 13.(Sheet 10)120

T.O. GR1F16CJ1

Cockpit Arrangement C DF (Typical) PX IIIDFDFINSTRUMENT
PANEL

LEFT
AUXILIARYCONSOLERIGHT
AUXILIARYCONSOLESTICK

THROTTLEDFRIGHTCONSOLELEFTCONSOLEDF

GR1F-16CJ-1-1006X37

Figure 13. (Sheet 11)

Change 1121

T.O. GR1F16CJ1

Cockpit Arrangement C DF (Typical) PX III
LEFT CONSOLE

24
23
25

1
2

22

21

4

19 20

7

5

18

16 17

14

15

8

13

10

12

3

6

9

DF
11
1. UHF Backup Control Panel
2. AUDIO 1 Control Panel
3. AUDIO 2 Control Panel
4. ECM Control Panel
5. AVTR Control Panel
6. EXT LIGHTING Control Panel
7. JHMCS Connection
8. MANUAL TRIM Panel
9. G-Suit Hose Connection
10. ANTI G TEST Button
11. DF STICK CONTROL Switch
12. Stowage
13. TEST Switch Panel
14. DEFOG Lever
15. FLT CONTROL Panel
16. Fuel Control Panel
17. IFF Panel
18. CANOPY JETTISON T-Handle
19. EPU Control Panel
20. ELEC Control Panel
21. Throttle FRICTION Control
22. ENG & JET START Control Panel
23. MANUAL PITCH Override Switch
24. CHAFF/FLARE Dispense Button
25. Throttle
GR1F-16CJ-1-1007X37

Figure 13. (Sheet 12)122

T.O. GR1F16CJ1

Cockpit Arrangement C DF (Typical) PX III
LEFT AUXILIARY CONSOLERIGHT AUXILIARY CONSOLE

DF

18

19

1

2

3

4
5

17

1

6

16
15
12

14
13

7

10

4

3

8

11

2

5

6
7

9

1. EMER STORES JETTISON Button (Covered)
2. WHEELS Down Lights (Green)
3. HOOK Switch (Lever Lock)
4. ANTI-SKID Switch
5. DN LOCK REL Button
6. LG Handle Down Permission Button
7. LG Handle
8. LG Handle Warning Light (Red)
9. LANDING TAXI LIGHTS Switch
10. CMDS Control Panel
11. HMCS Panel
12. RWR Aux Control Panel
13. ALT GEAR Handle
14. ALT GEAR Reset Button
15. SPEED BRAKE Position Indicator
16. STORES CONFIG Switch
17. HORN SILENCER Button
18. GND JETT ENABLE Switch (Lever Lock)
19. BRAKES Channel Switch

89

1. Magnetic Compass
2. FUEL Quantity Indicator
3. Pilot Fault List Display
4. System A HYD PRESS Indicator
5. System B HYD PRESS Indicator
6. Caution Light Panel
7. EPU FUEL Quantity Indicator
8. Cockpit Pressure Altimeter
9. Clock

GR1F-16CJ-1-1008X37

Figure 13. (Sheet 13)

123

T.O. GR1F16CJ1

Cockpit Arrangement C DF (Typical) PX III
INSTRUMENT PANEL

DF

1. HUD Combiner Glass
2. AOA Indexer
3. AR Status/NWS Indicator412
4. Integrated Control Panel (ICP)
5. Standby Attitude Indicator
6. FUEL FLOW Indicator40
7. Data Entry Display
39
8. ENG FIRE and ENGINE Warning
38
Lights (Red)37
35
9. HYD/OIL PRESS Warning Light (Red)36
10. FLCS and DBU ON Warning
34
Lights (Red)33
28
11. TO/LDG CONFIG Warning Light (Red)32
12. CANOPY and OXY LOW Warning Lights (Red)31
13. Right MFD
22
14. OIL Pressure Indicator29
15. NOZ POS Indicator30
16. RPM Indicator
2330. Autopilot ROLL Switch
17. FTIT Indicator
31. ADV MODE Switch
18. Vertical Velocity Indicator
32. MASTER ARM Switch
21
19. MRK BCN Light
33. LASER ARM Switch
20. FUEL QTY SEL Panel
34. IFF IDENT Button
21. Rudder PEDAL ADJ Knob
35. DRAG CHUTE Switch
22. AOA Indicator
36. RWR Prime Control Panel
23. INSTR MODE Select Panel
37. Azimuth Display/Indicator
24. Airspeed/Mach Indicator
38. F-ACK Button
25. Attitude Director Indicator
39. TF FAIL Warning Light (Red)26. Horizontal Situation Indicator
40. MASTER CAUTION Light (Amber)
27. Altimeter
41. ECM Light (Green), DF ECM (Green)
28. Left MFD
and OVRD (Amber) Light
29. Autopilot PITCH Switch

Figure 13. (Sheet 14)124Change 1

3

7
6 8
9

1
4

24

5
27

25
26 20

13
18
19

10
11

12
14
15
16
17

GR1F-16CJ-1-1009X37

T.O. GR1F16CJ1

Cockpit Arrangement C DF (Typical) PX III
RIGHT CONSOLE

17
16
DF
18

4

3

5

6

15
14

19

1
2

20
10

13

12

7
9 8
11

DF

1. SNSR PWR Control Panel
2. HUD Control Panel
3. Interior LIGHTING Control Panel
4. NUCLEAR CONSENT Switch (Guarded)
5. PLAIN Cipher Switch
6. AIR COND Control Panel
7. ENGINE ANTI-ICE Switch
8. Utility Light
9. ANT SEL Panel
10. AVIONICS POWER Panel
11. OBOGS MONITOR Panel
12. UDTU
13. Stowage
14. Oxygen/Communication Hookup
15. VOICE MESSAGE Switch
16. ZEROIZE Switch
17. Stick
18. DF NWS Control Button/Indicator (Green)
19. SEAT ADJ Switch
20. OXYGEN Panel
GR1F-16CJ-1-1010X37

Figure 13. (Sheet 15)

Change 1125

T.O. GR1F16CJ1

Cockpit Arrangement DR (Typical) PX III

INSTRUMENT
PANEL

LEFT
AUXILIARYCONSOLERIGHT
AUXILIARYCONSOLESTICK

THROTTLELEFT
CONSOLERIGHTCONSOLE

GR1F-16CJ-1-1011X37

Figure 13. (Sheet 16)126

T.O. GR1F16CJ1

Cockpit Arrangement DR (Typical) PX III
LEFT CONSOLE

7
9
8

1
2

6

10
5
11

4

3

1. AUDIO 1 Control Panel
2. AUDIO 2 Control Panel
3. G-Suit Hose Connection
4. ANTI G TEST Button
5. MAL & IND LTS Test Button
6. FUEL MASTER Switch
7. MANUAL PITCH Override Switch
8. ENG CONT Switch
9. Throttle
10. CANOPY JETTISON T-Handle
11. Stowage

GR1F-16CJ-1-1012X37

Figure 13. (Sheet 17)

127

T.O. GR1F16CJ1

Cockpit Arrangement DR (Typical) PX III
LEFT AUXILIARY CONSOLE

1

15

2

3

RIGHT AUXILIARY CONSOLE

4
7

14

5

12
10

6
1

9

11
8

6

13

3
4

1. EMER STORES JETTISON Button (Covered)
2. WHEELS Down Lights (Green)
3. HOOK Switch (Lever Lock)
4. DN LOCK REL Button
5. LG Handle Down Permission Button
6. LG Handle Warning Light (Red)
7. LG Handle
8. ALT GEAR Handle
9. ALT GEAR Reset Button
10. Blank Panel
11. SPEED BRAKE Position Indicator
12. ALT FLAPS Switch (Lever Lock)
13. HORN SILENCER Button
14. GND JETT ENABLE Switch (Lever Lock)
15. DRAG CHUTE Switch

Figure 13. (Sheet 18)128

2

5

1. Clock
2. FUEL Quantity Indicator
3. System B HYD PRESS Indicator
4. System A HYD PRESS Indicator
5. EJECTION MODE SELECT Handle
6. Caution Light Panel
GR1F-16CJ-1-1013X37

T.O. GR1F16CJ1

Cockpit Arrangement DR (Typical) PX III
INSTRUMENT PANEL

1. MASTER CAUTION Light (Amber) (NVIS)
40 41
38
2. ENG FIRE and ENGINE Warning
2 3
46
Lights (Red)37
39
3. HYD/OIL PRESS Warning Light (Red)1
4. FLCS and DBU ON Warning Lights (Red)5
5. AR Status/NWS Indicator
34
7 9
6. TO/LDG CONFIG Warning Light (Red)33
12
7. CANOPY and OXY LOW Warning Lights (Red)13
8
8. Right MFD (NVIS)
32
36
10
11
9. VIDEO SEL Switch
1435
10. Standby Attitude Indicator15
27 19 18
11. FUEL FLOW Indicator31
12. OIL Pressure Indicator
2117
13. FAULT ACK Button
28
20
16
14. NOZ POS Indicator
30 29 26
22
15. RPM Indicator
2329. MASTER MODE Caution Lights (NVIS) 25
16. FTIT Indicator
30. Data Entry Display
17. Pilot Fault List Display
24
31. Integrated Keyboard Panel (IKP) (NVIS)
18. Accelerometer
32. IFF IDENT Button
19. Altimeter
33. Stick Indicator
20. Vertical Velocity Indicator
34. Blank
21. Attitude Director Indicator
35. Azimuth Display/Indicator (NVIS)
22. Horizontal Situation Indicator
36. Left MFD (NVIS)
23. MRK BCN Light
37. AOA Indexer (NVIS)
24. Rudder PEDAL ADJ Knob
38. TF FAIL Warning Light (Red)
25. INSTR MODE Select Panel
39. Handgrip
26. AOA Indicator
40. ECM and OVRD Lights (Amber)
27. Airspeed/Mach Indicator
41. Aft Seat HUD Monitor (NVIS)
28. ARMT CONSENT Switch (Guarded)
GR1F-16CJ-1-1014X37

Figure 13. (Sheet 19)

129

T.O. GR1F16CJ1

Cockpit Arrangement DR (Typical) PX III
RIGHT CONSOLE

7
9

6

8

1

5

2
4

3

1. Interior LIGHTING Control Panel
2. OXYGEN REGULATOR Panel
3. Utility Light
4. Oxygen/Communications Hookup
5. Stowage
6. NWS Control Button/Indicator (Green)
7. Stick
8. SEAT ADJ Switch
9. STICK CONTROL/ASIU Panel

GR1F-16CJ-1-1015X37

Figure 13. (Sheet 20)

130

T.O. GR1F16CJ1

ENGINE PW 229

Main Fuel Control (MFC) PW 229

ENGINE FUEL SYSTEM PW 229

The MFC operates in both the PRI and SEC modes.
During PRI control, the MFC receives throttle inputs,
fuel from the main fuel pump, and electricalcommands from the DEEC. It controls main ignition,
start bleed strap position, main engine fuel flow, andrear compressor variable vane (RCVV) position. The
MFC also provides actuation pressure to thecompressor inlet variable vane (CIVV) control, theconvergent exhaust nozzle control (CENC), and both
the AB fuel control and AB pump controller.

Refer to figure 15. The engine fuel system delivers
the required fuel to the engine for combustion and foruse by the control system for scheduling the engine
variable geometry.

In SEC, the MFC receives throttle inputs, fuel fromthe main fuel pump, and static pressure and total
temperature signals from the fan inlet case. The MFC
controls main engine fuel flow, start bleed strap
position, RCVV's, and engine ignition.

ENGINE GENERAL DESCRIPTION PW 229
Refer to figure 14. The aircraft is powered by a single
F100PW229 afterburning turbofan engine. Maxi
mum thrust is approximately 29,000 pounds.

Afterburner (AB) Fuel Control PW 229

ENGINE CONTROL SYSTEM PW 229
The engine control system is composed of three major
components:the main fuel control (MFC), the AB
fuel control, and the digital electronic engine control(DEEC). The engine has two modes of opera
tion:primary (PRI) and secondary (SEC).

During primary operation, the AB fuel control
receives fuel from the AB fuel pump and electricalcommands from the DEEC. It provides AB ignition,
AB segment sequencing, and fuel flow to the AB
segments. During SEC control, AB fuel flow is
inhibited.

F100-PW-229 Engine
FAN MODULE

FAN
3 STAGES

CORE ENGINEMODULEFAN DRIVE
TURBINE
MODULEREAR COMPRESSOR
10 STAGESCOMBUSTIONCHAMBER
RCVV

FAN
DUCT

AB MODULE
AB SPRAY RINGSAND FLAME HOLDERCONVERGENTNOZZLE

DIVERGENTNOZZLEINLET
PRESSURE
PROBE
CIVV
ENGINEDIAGNOSTIC
UNIT

MAIN FUEL PUMPENGINE-DRIVEN GEARBOXAB LIGHT-OFF DETECTORHIGHPRESSURE TURBINEAB FUEL CONTROL
MAIN FUEL CONTROLDIGITAL ELECTRONIC ENGINE CONTROL
1F-16X-1-4004X

Figure 14.

131

T.O. GR1F16CJ1Engine Fuel/Control System Schematic
ENGINE F100-PW-229(Typical)LEGEND:
ELECTRICALMECHANICALENG BLEED AIR
FUEL (COOLING)

FUEL PRESSURIZATION AND
DUMP VALVEENGINE LUBE OIL TO
AND FROM OIL SYSTEM

MFP BOOST PRESSMFP PRESSURE
ENGINEDIAGNOSTIC
UNIT (EDU)

RCVV

CIVV

START
BLEED

SERVO PRESSURE
BOOST PUMP PRESS

FUEL/OIL
COOLERACTUATIONPRESSURE

ANTIICING
VALVE

ENGINE
SIGNALS

MAIN
FUEL
CONTROL(MFC)

DIGITAL
ELECTRONICENGINECONTROL(DEEC)

CONTROL AND
FEEDBACKENGINEAB FUEL
CONTROL(AFC)ENGINE CONTROLAND FEEDBACK
ENGINEFAN
DUCT

AB PUMPCONTROLLERAB FUEL
PUMP

LIGHT-OFF
DETECTORENGINEBLEED AIR(FOR ACTUATION)

AIRFRAME

ENGINE INLETCASE
ENGINEBLEEDAIR FORTURBINE
POWER

CIVV
ACTUATOR

CENC

OVBD

ENGINE BLEED AIRRCVVACTUATORAIRFRAME
SIGNALS

ENGINECONTROLANDFEEDBACKAIRFRAME SIGNALSAB FUEL FLOW

RTN
TO
RSVRS

MAIN
FUEL
PUMP

THROTTLE

FUEL FLOW
TRANSMITTER

FUEL LINESTRAINER

MAIN FUELSHUTOFF VALVE

FUEL/OIL HEATEXCHANGER
AB RESETCOCKPIT ANTI-ICE
ENGINE DATA REQUESTICE DETECTOR SIGNAL
MACH

ENGINE SIGNALSENGINE FAULTFTITNOZ POSOIL
RPMFROM FFPENGINE ELECTRONICCONTROL COOLING
FUEL SHUTOFF VALVE
GR1F-16CJ-1-1017X37

Figure 15.

132

T.O. GR1F16CJ1Digital Electronic Engine Control (DEEC) PW 229
The DEEC is an enginemounted, fuelcooled, solid
state digital full authority computer.
Control functions provided by the DEEC during PRI
operation are:
D Fuel flow scheduling.
D Fan speed control.
D Engine pressure ratio scheduling.
D Core compressor speed limiting.
D FTIT limiting.
D Minimumand maximum compressor dischargepressure limiting.

D Nozzle position control.
D RCVV scheduling.
D CIVV scheduling.
D Compressor start bleed system scheduling.
D AB segment sequencing and cancellation.
D Logic to detect and compensate for control system

malfunctions.

D Logic to provide automatic transfer to SEC in theevent of a critical PRI failure.

The DEEC closedloop idle control schedules MFC
fuel flow for three different idle levels. For functionaldescription, refer to ENGINE OPERATING CHARACTERISTICS PW 229 , this section.
The DEEC limits minimum engine rpm throughout theflight envelope to maintain stable operation. At highaltitude, low airspeed conditions, the DEEC protects
against engine stalls. During transonic and supersonicconditions, the DEEC limits minimum idle rpm as afunction of mach number (from CADC) to provide
sufficient engine airflow.
To minimize the possibility of stalls during AB operation
at high altitude and low airspeed, the DEEC commands
termination of segment 11 AB. At extremely highaltitude and low airspeed, the DEEC limits AB operation
to eight AB segments. When a stall is sensed, the DEEC
cancels the AB (if throttle is in AB range) and opens thenozzle until the stall clears. For subsequent AB
operation, the throttle must be retarded below AB before
AB can be reinitiated.

An engine overspeed or overtemperature condition
causes the DEEC to automatically transfer to SEC andilluminate the SEC caution light.
The DEEC receives power directly from the engine
alternator. In the event of an engine alternator or engine
gearbox failure indicated by rapid decrease to zero rpm
and illumination of the ENGINE warning light, theDEEC loses power and an automatic transfer to SEC
occurs.
Secondary Engine Control (SEC) PW 229
The SEC is a hydromechanical system which provides
engine control in the event of a DEEC malfunction. In
SEC, the CIVV's move to a fixed (cambered) position,
nozzle position is closed, the RCVV's are positioned by ahydromechanical control in the MFC, and AB operation
is inhibited. SEC is selected manually with the ENG
CONT switch or automatically by the DEEC. During
SEC operation, the SEC caution light illuminates.
Main Fuel Pump PW 229
The gearboxmounted main fuel pump provides
pressurized fuel to the MFC and boosts pressure to
the AB fuel pump.
Afterburner (AB) Fuel Pump PW 229
The AB fuel pump is driven by engine bleed air andprovides pressurized fuel to the AB. The pump operates
only during AB operation.
Compressor Inlet Variable Vane (CIVV)
Control PW 229
The CIVV control positions the CIVV's using MFC fuelpressure in response to an electrical signal from theDEEC. In SEC, the CIVV's are in a fixed (cambered)position.
Rear Compressor Variable Vanes (RCVV's) PW 229
The first three stages of the rear compressor are
equipped with variable geometry vanes. In PRI,RCVV's are controlled by the DEEC and are positioned
using pressurized fuel from the main fuel pump. In
SEC, the RCVV's are positioned by a hydromechanicalcontrol in the MFC as a direct function of throttle
position.

133

T.O. GR1F16CJ1
Compressor Bleed Air PW 229

LightOff Detector (LOD) PW 229

Lowpressure bleed air is directed from the bleed
strap into the fan duct to increase the compressor
stall margin during starting. Pressurized fuel fromthe main fuel pump is used to drive the start bleed
actuator. The bleed valve is scheduled as a function ofengine rpm by the DEEC when starting in PRI and as
a function of time and engine inlet pressure in SEC.

The engine incorporates an AB LOD, which, whencombined with the DEEC logic, provides AB nolightand blowout detection. When the LOD senses an ABnolight or blowout, the DEEC automaticallyterminates AB fuel flow. If the throttle is left in AB,
the DEEC reattempts AB lightoff up to three times.
If these attempts are unsuccessful, the throttle must
be retarded to MIL or below and then advanced into
AB for further AB attempts.

Highpressure bleed air is supplied to the EPU andengine nacelle ejectors. It is also used for engine inlet
antiicing, to drive the AB fuel pump, and to drive the
CENC motor.
Either lowpressure or highpressure air is provided
to the ECS depending on engine bleed pressure levels.
Pressurization and Dump Valve PW 229
A pressurization and dump valve is located in theengine fuel manifold line between the fuel/oil cooler
and fuel nozzles. It provides a minimum fuel pressure
for MFC operation at low rpm; the dump port iscapped so that fuel is not drained from the engine fuelmanifold when the throttle is retarded to OFF.
EXHAUST NOZZLE PW 229
The exhaust nozzle is variable and consists of two
sections. The divergent nozzle floats freely and moves
in conjunction with the convergent nozzle. The
convergent nozzle is controlled by the convergent
exhaust nozzle control.
Convergent Exhaust Nozzle Control(CENC) PW 229
The CENC is actuated by a highpressure bleed air
motor. The nozzle schedule is controlled by the DEEC
as a function of throttle input to the MFC. In PRI with
the LG handle down, the nozzle is greater than 80
percent open at IDLE (idle area reset). As the throttle
is advanced, the nozzle closes. With the LG handle up,
the nozzle is near minimum area except whenapproaching MIL or above. At MIL and above, theDEEC schedules the nozzle to control engine pressure
ratio as a function of fan speed. When the throttle isadvanced in the AB range, the DEEC schedules thenozzle open to compensate for increasing AB fuel flow.
In SEC, the nozzle is positioned to approximately zero
percent and AB operation is inhibited.

134ENGINE DIAGNOSTIC UNIT (EDU) PW 229
The EDU operates in conjunction with the DEEC toautomatically acquire and record diagnostic datawhenever the engine is operating. In the event that
the EDU detects one or more predetermined fault
codes associated with abnormal operation it willautomatically record approximately 4 seconds ofengine data (3 seconds before the event and one
second after the event). To manually acquire andrecord diagnostic data, place the AB RESET switch to
ENG DATA. Four seconds of engine data is recorded(3 seconds prior to switch movement and one second
after). The EDU can store one data set. The data
recorded by the ENG DATA switch overwrites theautomatically recorded data.
ENGINE OIL SYSTEM PW 229
The engine is equipped with a selfcontained oil
system to lubricate the engine and gearbox. Systempressure is nonregulated and varies with rpm and oiltemperature, and altitude.
Below approximately 35,000 feet MSL, oil pressure
should increase approximately 15 psi from IDLE to
MIL. At very high altitudes (50,000 feet), the oilpressure increase is approximately 5 psi from IDLE
to MIL. At all altitudes, however, a definite oilpressure increase should be evident when the rpm isincreased. Refer to SERVICING DIAGRAM, thissection for servicing/specifications information.
FUEL/OIL HOT Caution Light PW 229
The FUEL/OIL HOT caution light is located on thecaution light panel. The oil hot function of the light isinoperative.

T.O. GR1F16CJ1
ENGINE ANTIICE SYSTEM PW 229
The antiice system routes highpressure bleed air toand through the fixed fan inlet guide vanes, the
CIVV's, and the inlet pressure probe support cone toprevent ice formation. Additionally, the inlet pressureprobe is continuously heated electrically to prevent
ice formation. The system is controlled by the DEEC
and a threeposition ANTI ICE switch. The antiice
system can be activated manually by placing the
ANTI ICE switch to ON or automatically, if the ANTIICE switch is in AUTO and a sensor located in theinlet senses the accumulation of ice. Activation canalso occur if emergency dc bus No. 2 power is lost
(unless inhibited by the DEEC).

D OFF-Ice

detector, engine antiice system, andinlet strut heater are off.

INLET ICING Caution Light PW 229
The INLET ICING caution light located on thecaution light panel, illuminates when an ice
accumulation is detected by the inlet ice detector or if
a detection system failure occurs. The caution lightremains on for approximately 70 seconds (assuming
no additional ice accumulation). If more ice
accumulates, the caution light may remain on for alonger period of time or may cycle off and then on
again.

The inlet strut is electrically heated to prevent ice
buildup. This heater is also controlled by the ANTIICE switch for manual or automatic operation.

ENGINE AND ACCESSORY DRIVEGEARBOXES PW 229

The DEEC prevents antiice operation above 30,000
feet MSL and when engine inlet or bleed airtemperatures are high. In addition, a DEEC
malfunction may result in loss of bleed air for engine
antiicing.

Refer to figure 16. The engine gearbox drives themain fuel pump, the oil pump assembly, the engine
alternator, and the PTO shaft, which powers theaccessory drive gearbox (ADG).

Engine ANTI ICE Switch PW 229
The engine ANTI ICE switch is located on the rightconsole.
Functions are:
D ON-The inlet strut electrical heater turns on andthe engine antiice system is activated (if notinhibited by the DEEC). If ice accumulation isdetected, the INLET ICING caution light illumi
nates. The caution light remains on for approxi
mately 70 seconds (assuming that no additional ice
accumulation occurs). If ice reaccumulates before
the 70 second cycle expires, the caution lightremains on and the cycle repeats until icing
conditions no longer exist.

D AUTO-When an ice accumulation is detected, the

INLET ICING caution light illuminates, the inlet
strut electrical heater turns on, and the engine
antiice system activates (unless inhibited by theDEEC). The caution light, inlet strut electricalheater, and engine antiicing system remain on for
approximately 70 seconds (assuming that no
additional ice accumulation occurs). If ice reaccu
mulates before the 70 second cycle expires, thecaution light, inlet strut electrical heater, andengine antiicing system remain on and the cycle
repeats until icing conditions no longer exist.

The ADG powers the main generator through theconstantspeed drive (CSD), hydraulic system A andB pumps, standby generator, and FLCS PMG. The
JFS is also mounted on the ADG.
ENGINE ALTERNATOR PW 229
The engine alternator is driven by the engine gearbox
and provides sole power for the DEEC, engine and AB
ignition, inlet pressure probe heater, and the rpm
signal to the RPM indicator.
ENGINE IGNITION SYSTEM PW 229
The ignition system is powered by the engine
alternator and contains four igniter plugs (two for theengine and two for the AB). With the throttle at or
above IDLE and engine rpm at 12 percent or above,engine ignition is continuous. When the throttle is
moved into AB, AB ignition is activated by the DEEC
for up to 3 seconds or until the LOD detects an AB
light. In the event of an AB blowout or nolight
condition with the throttle left in AB, AB ignition is
automatically resequenced by the DEEC up to three
additional times. For subsequent AB ignition, thethrottle must be retarded to MIL or below and thenreturned to AB.

135

T.O. GR1F16CJ1
ENG & JET START CONTROL PANEL PW 229

JET FUEL STARTER (JFS) PW 229

Refer to figure 17. The ENG & JET START control
panel is located on the left console.

The JFS is a gas turbine which operates on aircraft
fuel and drives the engine through the ADG. The
JFS is connected by a clutch to the ADG and only
provides torque when required to maintain engine
rpm. If the ADG is not able to rotate (i.e., seizedengine), the JFS runs, but the clutch prevents it fromrotating the ADG. The JFS receives fuel at all times
regardless of the FUEL MASTER switch position.
The JFS is started by power from two brake/ JFS
accumulators used either singly or together. The
brake/JFS accumulators are charged automaticallyby hydraulic system B or manually by a hydraulichand pump located in the left wheel well. Automatic
recharging takes between 40 seconds (hot ambientconditions) and 60 seconds (cold ambient conditions).
The JFS is used to start the engine on the ground andto assist in engine airstart. Refer to JET FUELSTARTER LIMITS, Section V.

JFS Switch C

DF PW 229

Functions are:
D OFF - Normal switch position. The JFS can be shut

down at anytime by selecting OFF. The switch
returns to OFF automatically during a normal
ground start at 50 percent rpm.

D START 1-Vents one of the brake/JFS accumulators

to the hydraulic start motor.

D START 2-Vents both brake/JFS accumulators to

the hydraulic start motor.

JFS RUN Light C

DF PW 229

The green JFS RUN light illuminates within 30
seconds after initiating JFS start to indicate that the
JFS has attained governed speed.

Engine and Accessory Drive Gearboxes
ENGINE F100-PW-229

153

1. Standby Generator and FLCS PMG
2. Hydraulic System A Pump
3. Constant-Speed Drive
4. Main Generator
5. Engine Oil Pump
14
6. Tower Shaft (From Engine)
7. Main Fuel Pump
8. Engine Gearbox
9. Engine Alternator
10. JFS Exhaust Duct
11. Jet Fuel Starter
12. PTO Shaft
13. Hydraulic System B Pump
14. Accessory Drive Gearbox

4

8
9
11

13

10

12
GR1F-16CJ-1-1018X37

Figure 16.

136

7

6

2

T.O. GR1F16CJ1Once running, the JFS does not shut down until the
JFS switch is manually positioned to OFF.

JFS Operation PW 229
During a ground engine start, the brake/JFS
accumulators begin to recharge after the engine
accelerates through 12 percent rpm. As the engine
accelerates through 50 percent rpm, a sensor causes
the JFS to shut down automatically and the JFS RUN
light goes off.

ENGINE CONTROLS AND INDICATORS PW 229
Refer to figure 18. The engine instruments arelocated on the right side of the instrument panel.
Refer to ENGINE LIMITATIONS, Section V.
ENG CONT Switch PW 229

During inflight operation, the brake/JFS accumula
tors begin to recharge (provided system B hydraulicpressure is available) when the JFS reaches 70
percent of governed speed (34 seconds before the JFS
RUN light illuminates). If the JFS RUN light does notilluminate within 30 seconds or the JFS RUN lightgoes off once illuminated, the JFS START switch will
not reengage and the JFS cannot be restarted until
the JFS has spooled down. JFS spooldown takes
approximately 17 seconds from full governed speed.

The ENG CONT switch (guarded out of SEC) is
located on the left console. DR For ENG CONT switch
differences, refer to F16D AIRCRAFT, this section.
Functions are:
D C DF PRI-DEEC in operation (normal position).
D SEC-SEC operation. Transfer occurs when theswitch is moved to the SEC position.

ENG & JET START Control Panel (Typical)
1

ENGINE F100-PW-229

C

DF

3

DR

2

3

4
1. RUN Light (Green)
2. JFS Switch
3. ENG CONT Switch
4. AB RESET Switch
5. MAX POWER Switch

5

NOTE:
DR For ENG CONT switch,refer to F-16D AIRCRAFT,
this section.

GR1F-16CJ-1-1019X37

Figure 17.

137

T.O. GR1F16CJ1Engine Controls and Indicators (Typical)
ENGINE F100-PW-229

2

3

C

5

DF

6

7
8

4

9
10
11

ENGINE
FAULT

1

SEC

5

6

FUEL/OIL
HOT
INLET
ICING

12

OVERHEAT

13

7

8

DR

9
10

1

11

ENGINE
FAULT

2

SEC
FUEL/OIL
HOT
INLET
ICING
OVERHEAT
1. ENG CONT Switch
2. Throttle
3. C DF F-ACK Button,
DR FAULT ACK Button
4. Pilot Fault List Display

5. ENG FIRE and ENGINE Warning
Lights (Red)6. HYD/OIL PRESS Warning Light (Red)
7. FUEL FLOW Indicator
8. OIL Pressure Indicator

Figure 18.

138

4
3
12
9. NOZ POS Indicator
10. RPM Indicator
11. FTIT Indicator
12. Caution Lights (Amber)
13. ANTI ICE Switch
GR1F-16CJ-1-1020X37

T.O. GR1F16CJ1
AB RESET Switch C

DF PW 229

The AB RESET switch, located on the left console, isa threeposition toggle switch, springloaded to thecenter (NORM) position.
Functions are:
D AB RESET-This position is used to attempt toclear DEEC faults.

D NORM-Normal (deenergized) position.
D ENG DATA-This position may be used to record

engine data in the EDU.

ENGINE FAULT Caution Light PW 229
The ENGINE FAULT caution light, located on thecaution light panel, indicates that an engine PFL item
was detected. The caution light goes off when the fault
is acknowledged.
Pilot Fault List Display (PFLD) PW 229

expressed in percent from 0100. The indicator is
powered by battery bus No. 1.
NOZ POS Indicator PW 229
The NOZ POS indicator displays the position of the
CENC exhaust nozzle drive shafts which arecalibrated from 0 percent (closed) to 100 percent (fullyopen). The indicator accurately reflects exhaust
nozzle position in PRI and SEC unless both drive
shafts are failed. The indicator is powered byemergency ac bus No. 2.
FTIT Indicator PW 229
The FTIT indicator displays an average FTIT indegrees C. The indicator has a range of 200_1200_C
in major increments of 100_C and is powered bybattery bus No. 1.
FUEL FLOW Indicator PW 229

The PFLD, located on the C DF right auxiliary console,
DR instrument panel, displays engine PFL's. Refer to
WARNING, CAUTION, AND INDICATOR LIGHTS,
this section, for a description of the PFLD. Refer to
PILOT FAULT LIST-ENGINE, Section III, for a
description of engine PFL's.

The FUEL FLOW indicator is a digital indicator whichdisplays the total fuel flow to the engine, including AB,
in pph. The indicator has a range of 080,000 pph andis powered by emergency ac bus No. 1.

SEC Caution Light PW 229

The OIL pressure indicator displays engine oilpressure from 0100 psi and is powered by emergency
ac bus No. 2.

The SEC caution light, located on the caution light
panel, indicates that the engine is operating in SEC
or that main fuel pump pressure is low.
EEC Caution Light PW 229
The EEC caution light, located on the caution light
panel, is deactivated.
BUC Caution Light PW 229
The BUC caution light, located on the caution light
panel, is deactivated.
MAX POWER Switch C

DF PW 229

The MAX POWER switch, located on the left console,is inoperative.
RPM Indicator PW 229
The RPM indicator has a pointer display and the rpm
signal is supplied by the engine alternator. RPM is

OIL Pressure Indicator PW 229

HYD/OIL PRESS Warning Light PW 229
The HYD/OIL PRESS warning light, located on theedge of the right glareshield, serves as a monitor ofengine oil pressure and hydraulic system pressure.
For engine oil pressure, the warning light illumi
nates when oil pressure has been below approxi
mately 10 psi for 30 seconds (time delay minimizes
warning light illuminating during maneuvering).
The light goes out when oil pressure exceeds
approximately 20 psi. For hydraulic pressure, thewarning light illuminates when either A or B systempressure decreases below 1000 psi. The light goes out
when both system A and B pressures are above 1000
psi. During engine start, the light usually goes off
before reaching idle rpm; however, acceptable
operation is indicated if the warning light goes off
before exceeding 70 percent rpm and remains off
when the throttle is retarded to IDLE. The warning
light is powered by battery bus No. 1.

139

T.O. GR1F16CJ1
ENGINE Warning Light PW 229

D Ground idle - Provides the lowest level of idle thrust

The ENGINE warning light, located on the edge ofthe right glareshield, illuminates when RPM and
FTIT indicator signals indicate that an engine
overtemperature or flameout has occurred. Illumina
tion also occurs for an engine alternator failure andmay occur as a result of an RPM or FTIT indicator
failure. The warning light illuminates when the rpm
decreases to subidle (below 55 percent) or approxi
mately 2 seconds after FTIT indication exceeds
1100_C. The warning light goes off when thecondition that turned it on is eliminated. The warning
light is powered by battery bus No. 1.

D Flight idle - Flight idle provides inflight idle thrust

Throttle PW 229
Refer to figure 19. The engine is controlled by athrottle mounted above the left console with detents
at OFF, IDLE, MIL, and MAX AB. The throttle ismechanically connected to the MFC. The OFF
position terminates engine ignition and fuel flow. The
IDLE position commands minimum thrust and isused for all ground starts. From IDLE to MIL, thethrottle controls the output of the engine. Forward of
the MIL position, the throttle modulates theoperation of the AB (through 11 segments) while
maintaining constant basic engine operation.
C DF The throttle must be rotated outboard to allow
advancement from OFF to IDLE and from MIL to AB.
Retarding the throttle from AB to MIL automaticallyrotates the throttle. At IDLE, a cutoff release at thebase of the throttle must be actuated to allow thethrottle to be rotated outboard and retarded to OFF.
DR For throttle differences, refer to F16D AIR
CRAFT, this section.

A single white reflective stripe is located C DF on
both the upper surface of the throttle foot and on the
sidewall fairing, DR on both the lower throttle radius
next to the console and on the panel outboard of thethrottle radius. Alignment of the two stripes aids in
identifying the IDLE position.
Six switches are located on the throttle. C DF A
throttle friction control is located inboard at the baseof the throttle. DR The throttles are mechanicallylinked together.
ENGINE OPERATING CHARACTERISTICS PW 229
Engine General PW 229
Idle functions provided by the DEEC closedloop idle
control during PRI operation are:140

while maintaining adequate stall margin. The
nozzle opens to greater than 80 percent and engine
rpm is 6577 percent. Ground idle is activated with
the LG handle in DN and the throttle at or near
IDLE.
when the LG handle is in UP and the throttle is ator near IDLE. The nozzle is open to 020 percent andthe thrust is approximately 700 pounds higher thanground idle.

D Transient idle - Transient idle rpm provides for

rapid thrust response after rapidly retarding thethrottle to IDLE and then advancing within 20
seconds.

At MIL, the DEEC controls fan speed and enginepressure ratio to maintain consistent thrust. RPM
and FTIT vary as a function of flight conditions.
Following engine ground start, whenever the LG
handle is DN, and for 3 minutes after LG handle isplaced UP, the DEEC may position the RCVV's moreclosed for increased stall margin. Positioning the
RCVV's more closed results in up to 2 percent higher
engine rpm at MIL and above. Three minutes after
placing the LG handle UP, engine rpm may decrease
up to 2 percent.
Ground Operations PW 229
Since the DEEC maintains constant idle thrust, rpm
varies with temperature and pressure altitude(higher temperature or pressure altitude results inhigher rpm).
NonAB Operation in Flight PW 229
Regardless of temperature, NOZ POS indicator
indication should not exceed 20 percent at MIL.
Engine operation is continually optimized as flight
conditions change. This is evident by slight changes
in the NOZ POS, RPM, and FTIT indicator
indications.
Idle rpm is scheduled as a function of mach number(from CADC), altitude, temperature, throttle move
ment, and time. At altitudes below approximately
30,000 feet MSL, idle rpm is 7080 percent. As
altitude increases, idle rpm increases to provide theengine sufficient stall margin during throttle
transients.

T.O. GR1F16CJ1

Throttle (Typical)
ENGINE F100-PW-229

13
12

1

3

2

4

13

5
9

6

8
7

10

1. C DF Throttle Cutoff Release
2. UHF VHF (Data Link IN OUT) Transmit Switch (4-Way,
Momentary Rocker)
3. MAN RNG/UNCAGE Knob/Switch (Rotate, Depress)
4. ANT ELEV Knob (Rotate, Center Detent)
5. DOG FIGHT Switch (3-Position, Slide)
6. SPD BRK Switch (3-Position, Aft Momentary)
7. RDR CURSOR/ENABLE Switch (Depress Multidirectional)
8. Hands-on Blackout (HOBO) Paddle Switch
9. Throttle Foot
10. Throttle
11. C DF Throttle Friction Control
12. C DF IDLE Stripe
13. C DF Throttle Stripe

11

IDLE

AB

OFF
MIL (ADJUSTABLE)

MAX ABNOTE:
DR For throttle differences, refer to THROTTLE and F-16D AIRCRAFT, this section.

GR1F-16CJ-1-1021X37

Figure 19.

141

T.O. GR1F16CJ1
At 1.4 mach and above, the minimum thrust level is
MIL even though the throttle may be retarded below
MIL. Typically, the minimum thrust level increases
from idle to MIL between 0.981.4 mach. All of theminimum operating level features are deactivatedduring SEC operation.
After a rapid throttle movement to IDLE, engine rpm
initially decreases to a level above flight idle
(transient idle). Transient idle rpm provides for rapid
thrust response if the throttle is advanced. If thethrottle is not advanced within 20 seconds, engine
rpm then slowly decreases to flight idle. As altitude
increases, the difference between transient idle rpm
and flight idle rpm decreases.
A low frequency engine vibration may be sensed inflight or on the ground primarily at or near idle, but
may also occur at higher thrust settings. The
vibration has no adverse effect on engine or aircraft
structure and should disappear if engine rpm is either
increased or decreased. Vibrations that change in
intensity with throttle movement and are present
across the throttle/rpm range may indicate apotential engine malfunction.
AB Operation in Flight PW 229
Refer to figure 110. The DEEC monitors AB
operation and takes appropriate action to prevent
engine stalls. In AB, the DEEC provides thefollowing:
D Fastacceleration capability:The AB has nolimitations. Near sea level, AB operation occurs
immediately after AB is selected. At high altitude,
a higher fan speed must be attained prior to AB
operation. For example, during an IDLEtoMAX
AB throttle transient at low altitude, the AB lights
immediately when AB is selected and sequencing
begins just prior to attaining MIL thrust rpm.

D AB

segment sequencing limiting:When AB isselected at extremely high altitudes and low
airspeeds, the DEEC automatically schedules AB
operation. As the airspeed increases or the altitudedecreases, automatic AB sequencing takes place if
the AB request is greater than the actual AB
operation.

142Change 1

D AB recycle capability:The DEEC, in conjunction

with the LOD, provides automatic AB recycle
capability in the event of an AB blowout or nolight
condition (if the throttle is left in AB). In that event,the DEEC automatically resets the control system
to MIL, performs a control system check, andreattempts to light the AB up to three additionaltimes before returning the engine to MIL. If the
LOD is failed, the DEEC attempts one AB relightusing a duct pressure signal to verify AB lightoff. No
caution lights result from unsuccessful AB recycles.
Additional AB attempts can be made by moving thethrottle to MIL or below and then back into AB.

SEC Operation PW 229
The engine transfers to SEC when the ENG CONT
switch is manually switched to SEC. To minimize rpm
and thrust changes during manual transfers, thethrottle should be placed to the midrange position.
Transfer to SEC also occurs automatically if the
DEEC senses a major engine control system
malfunction or if loss of electrical power to the DEEC
occurs.
When the engine transfers to SEC, the SEC caution
light illuminates and AB operation is inhibited. RPM
and FTIT may increase or decrease depending on
flight conditions and on the engine malfunction.
If a transfer to SEC occurs while in AB, the nozzle
closes and AB operation is automatically cancelled. If
a transfer to SEC occurs during supersonic operation,
the throttle should be maintained at MIL or above
until the aircraft is subsonic.
While subsonic in SEC, throttle movement isunrestricted below 40,000 feet MSL. The throttle may
be moved in the AB range; however, the AB isinhibited. Refer to ENGINE - OPERATIONALENVELOPE, Section V for transfer and throttle
movement restrictions.
SEC provides 7080 percent of normal MIL thrust.
This level provides a measure of protection against
exceeding engine operating limits and provides
sufficient thrust for safe flight operations. SEC idle
thrust is approximately twice that in PRI with anormal nozzle during landing approach and ground
operations because the nozzle is closed.

T.O. GR1F16CJ1

AB Envelope

LightOff
ENGINE F100-PW-229

NOTES:
Throttle movement is unrestricted throughout the aircraft flight envelope.
Selecting AB above 45,000 feet MSL and less than 140 knots may result in delayed lights or recycles.
Region 1

Unlimited 11 segment AB operation.

Region 2

AB segments 1 through 10 available.

Region 3

AB segments 1 through 8 available.

Region 4

AB inhibited.
70

REGION 2
REGION 460

REGION 3

1000 FEET40

ALTITUDE

5030

REGION 1

20

10

0
0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

2.0

2.2

MACH NUMBER

1F-16X-1-4006X

Figure 110.

143

T.O. GR1F16CJ1

ENGINE GE 129

Digital Electronic Control (DEC) GE 129

GENERAL DESCRIPTION GE 129
Refer to figure 111. The aircraft is powered by asingle F110GE129 afterburning turbofan engine.
Maximum thrust is approximately 29,500 pounds.
ENGINE FUEL/CONTROL SYSTEM GE 129
Refer to figure 112. The engine fuel/control systemdelivers the required fuel to the engine for
combustion and for use by the control system forscheduling the engine variable geometry. The control
system is primarily composed of three major
components:the digital electronic control (DEC),
the afterburner fuel control (AFC), and the main
engine control (MEC). The engine has two pilot
selectable modes of operation:primary (PRI) andsecondary (SEC). In addition, there are two modes ofoperation between PRI and SEC which are notselectable by the pilot:hybrid VSV (HYB VSV) andhybrid (HYB).

The DEC is the critical component of the primary(PRI) engine control. The DEC is an enginemounted,
fuelcooled solidstate digital computer which con
trols both the main engine and the AB.
Afterburner Fuel Control (AFC) GE 129
The AFC is a fueloperated electrohydromechanicalcontrol which regulates fuel flow to the AB in
conjunction with the DEC.
Main Engine Control (MEC) GE 129
The MEC is a fueloperated hydromechanical control
which provides various control functions in all control
modes.
Primary (PRI) Engine Operation GE 129
PRI provides unrestricted engine operation throughout the entire flight envelope.

F110-GE-129 Engine
FAN MODULE

CORE ENGINEMODULEFAN DRIVE
TURBINE
MODULEREAR COMPRESSOR
9 STAGESFAN
3 STAGES

INLET
GUIDE
VANES

FAN
DUCT

AB MODULE
CONVERGENTNOZZLECOMBUSTIONCHAMBERAB FLAME DETECTORHIGH PRESSURE TURBINEAB FUEL CONTROL (LOWER RIGHT SIDE)
MAIN FUEL PUMPENGINE-DRIVENGEARBOX

DEC (LOWER RIGHT SIDE)

MAIN ENGINE CONTROL

Figure 111.

144

DIVERGENTNOZZLE

1F-16X-1-4005X

T.O. GR1F16CJ1Engine Fuel/Control System Schematic
ENGINE F110-GE-129(Typical)LEGEND:
ELECTRICALMECHANICAL
FUEL
FUEL (COOLING)
HYDRAULIC

FUEL/OIL/HYDRAULICCOOLERENGINE HYDRAULICPUMPNOZZLE TORQUE
MOTOR SIGNAL

MODE SELECT
SIGNALS
MEC THROTTLE
PRI/SEC SELECT
SIGNALENGINE CONTROLENGINEFEEDBACKSIGNALS

AB FLOW DEMANDAB FEEDBACK
DIGITAL
ELECTRONICCONTROL(DEC)

AB PUMP ON-OFFPYROMETER

MAIN
FUEL
PUMP

AB FUEL
CONTROL(AFC)

AB PUMP ON-OFF

MAIN
ENGINECONTROL(MEC)

PRI/SEC SELECT SIGNALSERVO PRESSURENOZ CONTVSV POWER

FAN IGV POWER

FAN DISCH TEMPAB FUEL FLOWAB FUELPUMP

AB PUMP SERVO PRESSURE
ENGINE FUEL
BOOST PUMPENGINE
AIRFRAME

THROTTLE

RTN
TO
RSVRS
FUEL FLOW
TRANSMITTERENGINE SIGNALSENGINE FAULTFTITNOZ POSOIL
RPM

MAIN FUELSHUTOFF VALVECOCKPIT ANTI-ICE
ENGINE DATA REQUESTICE DETECTOR SIGNAL
FUEL/OIL HEATMACHEXCHANGERFROM FFPAIRFRAME SIGNALSENGINE ELECTRONICCONTROL COOLING
FUEL SHUTOFF VALVE
GR1F-16CJ-1-0017X37

Figure 112.

145

T.O. GR1F16CJ1Control functions provided by the DEC during PRI
operation are:
D Fan speed control.
D Core speed limiting.
D Acceleration and deceleration fuel flow scheduling.
D Turbine blade temperature limiting.

During transonic and supersonic flight, with thethrottle retarded below MIL, the DEC limits
minimum engine operation as a function of machnumber from the central air data computer (CADC)
to prevent inlet buzz and possible engine stall. Whenretarding the throttle to IDLE above 1.4 mach, rpm
may decrease up to 15 percent from MIL rpm. RPM
then decreases with mach number until approxi
mately 1.1 mach, at which time the engine
decelerates to normal flight idle rpm.

D AB fuel flow scheduling.
D Nozzle control to provide fan stall margin.
D Minimumand maximum compressor dischargepressure limiting.

D Scheduling of inlet guide vane (IGV) position.
D Resetting of the compressor variable stator vanes(VSV) for increased stall protection.

D Ignition logic for starting and automatic relightsequencing in both the engine and AB.

D Logic to automatically select HYB or HYB VSV ortransfer to SEC for certain PRI failures.

D Compressor variable stator vane scheduling.

Control functions provided by the MEC during PRI
operation are:Hybrid (HYB) Engine Operation GE 129
HYB is activated when the DEC detects certain
failures. In HYB, the MEC provides main engine fuelflow scheduling and VSV control.
Control functions by the DEC during HYB operationare:
D AB fuel flow scheduling.
D Nozzle control to provide fan stall margin.
D Scheduling of inlet guide vane (IGV) position.
D Logic to automatically transfer to SEC if HYB fails.

Control functions provided by the MEC during HYB
operation are:

D Main engine fuel flow scheduling and metering.

D Main engine fuel flow scheduling and metering.

D Engine

D Compressor variable stator vane (VSV) scheduling.

overspeed protection (113 percent rpm
overspeed fuel shutoff valve).

D Positive fuel cutoff.
D Compressor VSV scheduling (HYB VSV) for certain

PRI failures.

When operating in PRI, main engine fuel flow iscontrolled by the DEC. The MEC fuel flow control
feature is in standby mode.
The nozzle is controlled by signals from the DEC tothe engine hydraulic pump which positions four
nozzle actuators in order to maintain fan stall margin
while providing the requested level of thrust.

D Engineoverspeed protection (113 percent rpm
overspeed fuel shutoff valve).

D Positive fuel cutoff.

During HYB operation:
D ENG HYB MODE PFL is displayed.
D Turbine blade temperature limiting is not provided.
D VSV reset is not active.
D Maximum fan speed is automatically limited.

Fan inlet guide vane (IGV) positioning is controlled
by the DEC in accordance with the IGV schedule.

D MIL thrust is 90100 percent of that provided in

High energy and AB ignition are controlled by theDEC.

D Supersonic idle - Lockup is not active.

146

PRI.

T.O. GR1F16CJ1Secondary Engine Control (SEC)
Operation GE 129
SEC is activated by either manually placing the ENG
CONT switch to SEC or as a result of automatictransfer when the DEC detects certain failures. In
SEC, the MEC provides fuel flow scheduling in
addition to the functions it provides in PRI operation.
During SEC operation:
D The nozzle is closed.

Inlet Guide Vanes (IGV's) GE 129
Each IGV is an airfoil which is divided into two
sections. The forward portion of the inlet guide vane
is fixed which provides structural support. The aft
portion of the inlet guide vane is a variable angle
flap which controls the angle at which air enters thefan. This both improves fan efficiency and increases
the stall margin.
Variable Stator Vanes (VSV's) GE 129

D SEC caution light is illuminated.

The compressor VSV system controls the angle ofthe core inlet guide vanes and the first three stagesof core variable stator vanes. Positioning is afunction of engine rpm. By varying the vane
position, the system automatically changes theeffective angle at which the airflow enters thecompressor rotor blades, thereby maintaining
satisfactory airflow and optimum compressor
performance throughout the entire flight envelope.
For increased stall protection, the VSV's are reset
slightly closed from their normal position after athrottle snap to IDLE. The reset position is
maintained for 2 minutes after which the VSV's
return to their normal schedule, resulting in an rpm
drop of approximately 2 percent.

D In

Compressor Bleed Air GE 129

D AB operation is inhibited (fuel and ignition).
D Turbine blade temperature limiting is not provided.
D High energy ignition is continuously energized.
D IGV's are in a fixed, fully closed position.
D VSV reset is not active.
D Maximum fan speed is automatically limited.

flight, MIL thrust is 7095 percent of thatprovided at PRI MIL.

D Supersonic idle - Lockup is not active.
D Idle thrust is higher than that in PRI because thenozzle is closed.

Engine Fuel Boost Pump GE 129
The gearboxmounted engine fuel boost pumpprovides pressurized fuel to the main fuel pump andAB fuel pump.
Main Fuel Pump GE 129

Bleed air is extracted from two separate stages in
the compressor for engine and airframe use.
Lowpressure (fifth stage) air is used for turbine
cooling and the engine antiice system. Air for
airframe use is taken from both the lowandhighpressure (ninth stage) compressor sections.
Lowpressure bleed air is used for the ECS unlessthe pressure is insufficient, in which casehighpressure bleed air is used. Highpressure
bleed air is used for the nacelle ejectors and is also
used to power the EPU.
EXHAUST NOZZLE GE 129

Afterburner (AB) Fuel Pump GE 129

The exhaust nozzle is a variable area convergent/di
vergent, semifloating type with mechanically linked
primary and secondary flaps and seals. Nozzle area
modulation is accomplished by four hydraulicactuators which provide synchronous actuation. The
nozzle actuators are operated by the engine hydraulicpump using engine oil as hydraulic fluid, and respondto electrical inputs from the DEC.

The gearboxmounted AB fuel pump receives fuelfrom the engine fuel boost pump. It provides
additional pressure and supplies fuel to the AB fuelcontrol.

The primary functions of the nozzle system are to
maintain fan stall margin by varying the nozzle area
and to control the engine thrust for optimum
performance through the entire flight envelope.

The geartype main fuel pump receives pressurizedfuel from the engine fuel boost pump. It provides
additional pressure and supplies the fuel to theMEC.

147

T.O. GR1F16CJ1
AB FLAME DETECTOR GE 129
The AB flame detector provides AB light/nolightinformation to the DEC for use in AB sequencing andautorelight functions during PRI operation.

ENGINE MONITORING SYSTEM (EMS) GE 129
The EMS is operative in all engine control modes(PRI, SEC, HYB VSV, and HYB).
The EMS is designed to perform engine diagnostics
and store engine fault data for postflight analysis.
The EMS consists of two primary components:thedigital electronic control (DEC) which is mounted on
the engine and an engine monitoring systemcomputer (EMSC) which is located in the leading edge
flap drive bay in the forward fuselage section.
In all engine control modes the EMSC receives datafrom both the aircraft and the DEC to perform engine
diagnostics and fault detection. Upon detection of a
fault, the EMSC automatically stores approximately
8 seconds of engine data (6 seconds before and 2
seconds after the fault). If the fault is an engine pilot
fault list (PFL) item, the ENGINE FAULT caution
light illuminates.
Event data can also be manually stored at anytime bymomentarily placing the AB RESET switch to ENG
DATA. This function may be especially useful in eventof unusual engine operation where EMS data was not
automatically recorded. This action has no effect ondata that was automatically recorded.
Data recorded by the EMS are parts life tracking
data, engine performance trending data, and event
data, including PFL's and MFL's. This data can beextracted from the EMS by maintenance personnel
after flight.

148

During HYB VSV, HYB, and SEC modes, the EMSC
continues the monitoring and reporting function for
faults associated with the particular mode ofoperation.
ENGINE OIL SYSTEM GE 129
The engine is equipped with a selfcontained, dry
sump, full pressure lubrication system which
provides filtered oil for lubricating and cooling theengine main shaft bearings, oil seals, gearboxes, andaccessories. It also provides oil to the engine
hydraulic pump for nozzle actuation. Engine oil level
decreasing below approximately 40 percent of normal
capacity for 15 seconds results in activation of an
ENG LUBE LOW PFL. Refer to SERVICINGDIAGRAM, this section for servicing/specifications
information.
FUEL/OIL HOT Caution Light GE 129
The FUEL/OIL HOT caution light, located on thecaution light panel, illuminates when the temperature of the oil becomes excessive (exceeds 300_F). The
caution light also illuminates as a function ofexcessive fuel temperature. Refer to FUEL SYSTEM,
this section.
ENGINE ANTIICE SYSTEM GE 129
The engine antiice system prevents the formation
and accumulation of ice on the front frame struts, theforward centerbody, and the IGV flaps. Lowpressure
bleed air is directed to a pressure regulatingantiicing valve. Valve actuation is accomplishedeither automatically (with the ANTI ICE switch in
AUTO position) by an ice detector in the engine inlet
or manually by the ANTI ICE switch. When theantiicing valve is open, lowpressure bleed air entersthe front frame at two separate locations. Bleed air isthen distributed radially to each of the front frame
struts, the IGV flaps, and the forward centerbody.
Automatic activation also occurs if emergency dc bus
No. 2 power is lost.
If engine antiicing is either manually or automati
cally demanded while operating at idle rpm, engine
rpm may increase as much as 3 percent when ambient
temperature is below 30°F.
The inlet strut is electrically heated to prevent ice
buildup. The heater is also controlled by the ANTIICE switch for manual or automatic operation.

T.O. GR1F16CJ1Engine ANTI ICE Switch GE 129

ENGINE AND ACCESSORY DRIVEGEARBOXES GE 129

The engine ANTI ICE switch is located on the rightconsole.

Refer to figure 113. The engine gearbox drives themain fuel pump, the engine fuel boost pump, the AB
fuel pump, the engine/scavenge pump, the engine
alternator, the engine hydraulic pump, the MEC,
and the PTO shaft, which powers the accessory
drive gearbox (ADG).

Functions are:
D ON - The inlet strut electrical heater turns on andthe engine antiice system is activated. If ice
accumulation is detected, the INLET ICING
caution light illuminates. The caution lightremains on for approximately 70 seconds (assum
ing that no additional ice accumulation occurs). If
ice reaccumulates before the 70second cycle
expires, the caution light remains on and the cycle
repeats until icing conditions no longer exist.

D AUTO - When an ice accumulation is detected,

the INLET ICING caution light illuminates, theinlet strut electrical heater turns on, and theengine antiice system activates. The caution
light, inlet strut electrical heater, and engine
antiicing system remain on for approximately 70
seconds (assuming that no additional ice accu
mulation occurs). If ice reaccumulates before the70second cycle expires, the caution light, inlet
strut electrical heater, and engine antiicing
system remain on and the cycle repeats until icing
conditions no longer exist.

D OFF-Ice

detector, engine antiice system, andinlet strut heater are off.

INLET ICING Caution Light GE 129
The INLET ICING caution light, located on thecaution light panel, illuminates when an ice
accumulation is detected by the inlet ice detector orif a detection system failure occurs. The caution
light remains on for approximately 70 seconds(assuming no additional ice accumulation). If more
ice accumulates, the caution light may remain on
for a longer period of time, or it may cycle off andthen on again.

The ADG powers the main generator through theconstantspeed drive (CSD), system A and B
hydraulic pumps, standby generator, and FLCS
PMG. The JFS is also mounted on the ADG.
ENGINE ALTERNATOR GE 129
The engine alternator is driven by the engine
gearbox and provides the rpm signal to the RPM
indicator and power for high energy and AB ignition
and the DEC.
ENGINE IGNITION SYSTEM GE 129
The ignition system contains three igniter plugs (twofor the engine and one for the AB). The engine igniters
are controlled by engine rpm and operate completely
independent of throttle position or throttle move
ment.
During engine start, high energy ignition, which is
powered by the engine alternator, is automaticallyturned on at approximately 10 percent rpm and is
automatically turned off at 59 percent rpm. The
ignition system also provides an automatic relightfeature which selects high energy ignition when theengine rpm deceleration rate exceeds 5 percent per
second, or when the engine rpm goes below 59 percent.
A low energy ignition system is used as a backup to
the high energy system. The low energy system isenergized if the engine decelerates through approxi
mately 55 percent rpm while airborne. The ignition
system is powered by emergency ac bus No. 1.
AB ignition is powered by the engine alternator andprovides energy to the AB spark igniter. AB ignition
is automatically controlled by the DEC.

149

T.O. GR1F16CJ1Engine and Accessory Drive Gearboxes
ENGINE F110-GE-129

6
1

7

9

8

2
5

3
4

10
11

1. Standby Generator and FLCS PMG
2. System A Hydraulic Pump
3. Constant-Speed Drive
4. Main Generator
5. AB Fuel Pump
6. Tower Shaft (From Engine)18
7. Fuel Boost Pump
8. Engine Alternator
179. Main Engine Control
10. Main Fuel Pump
14. JFS Exhaust Duct
11. Engine Gearbox
15. Jet Fuel Starter
12. Engine Lube/Scavenge Pump
16. PTO Shaft
13. Engine Hydraulic Pump

12
13
16

15

14

17. System B Hydraulic Pump
18. Accessory Drive Gearbox
GR1F-16CJ-1-0018X37

Figure 113.

JET FUEL STARTER (JFS) GE 129

ENG & JET START CONTROL PANEL GE 129

The JFS is a gas turbine which operates on aircraft
fuel and drives the engine through the ADG. The
JFS is connected by a clutch to the ADG and only
provides torque when required to maintain engine
rpm. If the ADG is not able to rotate (i.e., seizedengine), the JFS runs, but the clutch prevents it fromrotating the ADG. The JFS receives fuel at all times
regardless of the FUEL MASTER switch position.
The JFS is started by power from two brake/JFS
accumulators used either singly or together. The
brake/JFS accumulators are charged automaticallyby hydraulic system B or manually by a hydraulichand pump located in the left wheel well. Automatic
recharging takes between 40 seconds (hot ambientconditions) and 60 seconds (cold ambient condi
tions). The JFS is used to start the engine on the
ground and to assist in engine airstart. Refer to JET
FUEL STARTER LIMITS, Section V.

Refer to figure 114. The ENG & JET START control
panel is located on the left console.

150JFS Switch C

DF

GE 129Functions are:
D OFF - Normal switch position. The JFS can be shut

down at anytime by selecting OFF. The switch
returns to OFF automatically during a normal
ground start at approximately 55 percent rpm.

D START 1 - Vents one of the brake/JFS accumulators

to the hydraulic start motor.

D START 2 - Vents both brake/JFS accumulators to

the hydraulic start motor.

T.O. GR1F16CJ1

ENG & JET START Control Panel (Typical)
1

ENGINE F110-GE-129

C

DF

3

DR

2

3

4

1. RUN Light (Green)
2. JFS Switch
3. ENG CONT Switch
4. AB RESET Switch
5. MAX POWER Switch

NOTE:
DR For ENG CONT switch,refer to F-16D AIRCRAFT,
this section.

5

GR1F-16CJ-1-0019X37

Figure 114.

JFS RUN Light C

DF

GE 129The green JFS RUN light illuminates within 30
seconds after initiating JFS start to indicate that the
JFS has attained governed speed.
JFS Operation GE 129
During a ground engine start, the brake/JFS
accumulators begin to recharge after the engine
accelerates through 12 percent rpm. As the engine
accelerates through approximately 55 percent rpm, asensor causes the JFS to shut down automaticallyand the JFS RUN light goes off.
During inflight operation, the brake/JFS accumula
tors begin to recharge (provided system B hydraulicpressure is available) when the JFS reaches 70
percent of governed speed (34 seconds before the JFS
RUN light illuminates). If the JFS RUN light does notilluminate within 30 seconds or the JFS RUN lightgoes off once illuminated, the JFS START switch will
not reengage and the JFS cannot be restarted until

the JFS has spooled down. JFS spooldown takes
approximately 17 seconds from full governed speed.
Once running, the JFS does not shut down until the
JFS switch is manually positioned to OFF.
ENGINE CONTROLS AND INDICATORS GE 129
Refer to figure 115. The engine instruments arelocated on the right side of the instrument panel.
Refer to ENGINE LIMITATIONS, Section V.
ENG CONT Switch GE 129
The ENG CONT switch (guarded out of SEC) is
located on the left console. DR For ENG CONT switch
differences, refer to F16D AIRCRAFT, this section.
Functions are:
D C DF PRI - PRI operation (normal position).
D SEC - SEC operation. Transfer occurs when theswitch is moved to the SEC position.

151

T.O. GR1F16CJ1Engine Controls and Indicators (Typical)
ENGINE F110-GE-129

2

3

C

5

DF

6

7
8

4

9
10
11

ENGINE
FAULT

1

SEC

5

6

FUEL/OIL
HOT
INLET
ICING

12

OVERHEAT

13

7

DR

8

3

9
10

1
11

ENGINE
FAULT

2

SEC
FUEL/OIL
HOT
INLET
ICING
OVERHEAT
1. ENG CONT Switch
2. Throttle
3. C DF F-ACK Button,
DR FAULT ACK Button
4. Pilot Fault List Display

5. ENG FIRE and ENGINE Warning
Lights (Red)6. HYD/OIL PRESS Warning Light (Red)
7. FUEL FLOW Indicator
8. OIL Pressure Indicator

Figure 115.

152

4

12
9. NOZ POS Indicator
10. RPM Indicator
11. FTIT Indicator
12. Caution Lights (Amber)
13. ANTI ICE Switch
GR1F-16CJ-1-0020X37

T.O. GR1F16CJ1SEC Caution Light GE 129

RPM Indicator GE 129

The SEC caution light, located on the caution light
panel, indicates that the engine is operating in SEC.

The RPM indicator has a pointer display expressed inpercent rpm from 0110. The rpm signal is supplied by
the engine alternator. The indicator is powered bybattery bus No. 1.

EEC Caution Light GE 129
The EEC caution light, located on the caution light
panel, is deactivated.
BUC Caution Light GE 129
The BUC caution light, located on the caution light
panel, is deactivated.
MAX POWER Switch C

DF

GE 129NOZ POS Indicator GE 129
The NOZ POS indicator is a direct display of actual
nozzle position ranging from 0 percent (closed) to 100
percent (full open). The indicator is powered byemergency ac bus No. 2.
FTIT Indicator GE 129

The MAX POWER switch, located on the left console,is inoperative.

The FTIT indicator displays exhaust gas temperature(EGT) in degrees C. The indicator has a range of
200_1200_C in major increments of 100_C and is
powered by battery bus No. 1.

AB RESET Switch C

FUEL FLOW Indicator GE 129

DF

GE 129The AB RESET switch, located on the left console, isa threeposition toggle switch, springloaded to center(NORM) position.
Functions are:
D AB RESET - Inoperative position.
D NORM - Normal (deenergized) position.
D ENG DATA - This position is used to record engine

data in the EMSC. When ENG DATA is
momentarily selected, ENG 066 MFL is generated
and engine data is recorded for an 8second period.
Recorded data begins 6 seconds prior to selecting
ENG DATA.

ENGINE FAULT Caution Light GE 129
The ENGINE FAULT caution light, located on thecaution light panel, indicates that an engine PFL
item was detected. The caution light goes off when thefault is acknowledged.
Pilot Fault List Display (PFLD) GE 129
The PFLD, located on the C DF right auxiliary
console, DR instrument panel, displays engine PFL's.
Refer to WARNING, CAUTION, AND INDICATOR
LIGHTS, this section, for a description of the PFLD.
Refer to PILOT FAULT LIST - ENGINE, Section III,
for a description of engine PFL's.

The FUEL FLOW indicator is a digital indicatorwhich displays the total fuel flow to the engine,
including AB, in pph. The indicator has a range of
080,000 pph and is powered by emergency ac bus
No. 1.
OIL Pressure Indicator GE 129
The OIL pressure indicator displays engine oilpressure from 0100 psi and is powered by emergency
ac bus No. 2.
HYD/OIL PRESS Warning Light GE 129
The HYD/OIL PRESS warning light, located on theedge of the right glareshield, serves as a monitor ofengine oil pressure and hydraulic system pressure.
For engine oil pressure, the warning light illumi
nates when oil pressure has been below approxi
mately 10 psi for 30 seconds (time delay minimizes
warning light illuminating during maneuvering).
The light goes out when oil pressure exceeds
approximately 20 psi. For hydraulic pressure, thewarning light illuminates when either A or B systempressure decreases below 1000 psi. The light goes out
when both system A and B pressures are above 1000
psi. During engine start, the warning light usually
goes off before reaching idle rpm; however, acceptable
operation is indicated if the warning light goes off
before exceeding 70 percent rpm and remains off
when the throttle is retarded to IDLE. The warning
light is powered by battery bus No. 1.

153

T.O. GR1F16CJ1
ENGINE Warning Light GE 129
The ENGINE warning light, located on the edge ofthe right glareshield, illuminates when RPM and
FTIT indicator signals indicate that an engine
overtemperature or flameout has occurred. Illumina
tion also occurs for an engine alternator failure andmay occur as a result of an RPM or FTIT indicator
failure. The warning light illuminates when the rpm
decreases to subidle (below 60 percent) or approxi
mately 2 seconds after FTIT indication exceeds
1100_C. The warning light goes off when thecondition that turned it on is eliminated. The warning
light is powered by battery bus No. 1.
Throttle GE 129
Refer to figure 116. The engine is controlled by athrottle mounted above the left console with detents
at OFF, IDLE, MIL, and MAX AB. The OFF position
inhibits fuel flow. The IDLE position commands
minimum thrust and is used for all ground starts.
From IDLE to MIL, the throttle controls the outputof the engine. Forward of the MIL position, thethrottle modulates the operation of the AB while
maintaining constant basic engine operation.
C DF The throttle must be rotated outboard to allow
advancement from OFF to IDLE and from MIL to
AB. Retarding the throttle from AB to MIL
automatically rotates the throttle. At IDLE, a cutoff
release at the base of the throttle must be actuatedto allow the throttle to be rotated outboard andretarded to OFF. DR For throttle differences, refer toF16D AIRCRAFT, this section.

Six switches are located on the throttle. C DF A
throttle friction control is located inboard at the baseof the throttle. D The throttles are mechanicallylinked together.
ENGINE OPERATING CHARACTERISTICS GE 129
Ground Operations GE 129
During engine start, the SEC caution light remains
on until approximately 20 percent rpm. Since the
DEC maintains constant idle thrust and minimum
bleed air pressure, rpm varies with temperature andpressure altitude (higher temperature or pressure
altitude results in higher rpm). While VSV reset isactive, idle rpm is 23 percent higher than normal.
During cold weather operations below 30_F, the DEC
logic, when either commanded manually or automat
ically, increases antiice bleed air pressure to
improve system performance at idle rpm. Engine
rpm may increase as much as 3 percent duringantiice operations.
Engine vibration/rumble may be sensed in flight or on
the ground while accelerating or decelerating theengine, when transferring to SEC, or when stabilized
below MIL power in either PRI or SEC. The
vibration/rumble at a stabilized rpm has no adverse
effect on the engine or aircraft structure. The
vibration/rumble should disappear when the throttle
is advanced slightly (approximately 5 percent rpm
increase). Vibration that persists or becomesappreciably more intense when the throttle isadvanced may indicate a potential engine malfunc
tion.

Both electrical and mechanical throttle positions aretransmitted to the engine. The electrical input isused by the DEC as the primary throttle position
signal to compute various engine operating values.
The throttle is also mechanically connected to theengine via a cable to the MEC. This mechanicalthrottle position input is used for main fuel cutoff
and during SEC and HYB operation. It also provides
a backup electrical signal to the DEC during PRI
operation.

Engine operation is continually optimized as flight
conditions change. This is evident by slight changes
in the NOZ POS, RPM, and FTIT indicator
indications.

A single white reflective stripe C DF is located on
both the upper surface of the throttle foot and on the
sidewall fairing, DR on both the lower throttle radius
next to the console and on the panel outboard of thethrottle radius. Alignment of the two stripes aids in
identifying the IDLE position.

As altitude increases, idle rpm increases in order tomaintain the required bleed air pressure forsatisfactory ECS operation and stall margin for theengine. As a result, rpm and FTIT response during
throttle movement between low throttle settings issignificantly reduced.

154

NonAB Operation in Flight GE 129
Regardless of temperature, stabilized NOZ POS
indicator indications should not exceed 15 percent
open in MIL.

T.O. GR1F16CJ1

Throttle (Typical)
ENGINE F110-GE-129

12
11
1

2

3

4

12
5
8

6
7

9
10

1. C DF Throttle Cutoff Release
2. UHF VHF IFF Transmit Switch (4-Way, Momentary Rocker)
3. MAN RNG/UNCAGE Knob/Switch (Rotate, Depress)
4. ANT ELEV Knob (Rotate, Center Detent)
5. DOG FIGHT Switch (3-Position, Slide)
6. SPD BRK Switch (3-Position, Aft Momentary)
7. RDR CURSOR/ENABLE Switch (Depress, Multidirectional)
8. Throttle Foot9. Throttle
10. C DF Throttle Friction Control
11. C DF IDLE Stripe
12. C DF Throttle Stripe

OFF

IDLE

AB

MIL (ADJUSTABLE)

MAX ABNOTE:
DR For throttle differences, refer to THROTTLE and F-16D AIRCRAFT, this section.
GR1F-16CJ-1-0021X37

Figure 116.

155

T.O. GR1F16CJ1
At 1.4 mach and greater, the minimum thrust level isnear MIL even though the throttle may be retarded
below MIL. Typically, the minimum thrust level
decreases with mach number between 1.41.1 mach.
At IDLE and while decelerating through 1.1 mach,the engine decelerates to normal inflight idle thrust.
Reduced speed excursion (RSE) logic is activatedduring engine deceleration to idle speed above 0.6
mach. RSE results in a higher inflight idle rpm offset
by a greater nozzle opening than normal inflight idle.
Idle thrust changes from inflight idle to approachidle between 0.50.6 mach, resulting in an engine rpm
change of approximately 10 percent. Slightly moretime is required to accelerate the engine fromapproach idle thrust when airspeed is below 0.5
mach.
Idle thrust changes from approach idle to ground idle
between 8090 knots and results in reduced engine
rpm of approximately 2 percent. This change occurs
during the landing roll to achieve desired ground idle
thrust levels for taxi.
While VSV reset is active, idle rpm is 23 percent
higher than normal.
A high frequency vibration may be felt through the
cockpit floor, ejection seat, and/or rudder pedals as aresult of aircraft structural response to normal
engine operation. The vibration is most noticeablewith the aircraft in a clean configuration with
reduced fuel loads at lower airspeeds when the engine
is operating near MIL thrust or above. This vibration
has no adverse effect on the engine or aircraft.
AB Operation GE 129
Refer to figure 117 for AB operational characteris
tics. During AB operation, FTIT, rpm, and oilpressure vary with altitude and airspeed. NOZ POS
indications during minimum AB operation should be
between 717 percent open and for MAX AB
operation, between 4070 percent open. The engine
has a reduced AB region of operation (at highaltitudes and low airspeeds) and commands adecrease in AB fuel flow to prevent AB instabilities.
The NOZ POS indicator indicates approximately
3050 percent open during this reduced AB operation.

156

When AB operation is first initiated, the exhaust
nozzle preopens up to 10 percent more than MIL
exhaust nozzle position to increase stall margin
during AB lightoff. Fuel flow and exhaust nozzle
area are held at minimum AB levels until the flame
detector determines that lightoff (within 5 seconds(greater than 40_F) or 10 seconds (40_F or less) of AB
selection) has occurred. Once AB lightoff occurs, fuelflow and exhaust nozzle area increase to therequested AB level with a corresponding increase inthrust. If AB blowout occurs, the autorelight featureattempts to reinitiate AB without throttle movement.
During AB operation at or above 45,000 feet, a mild
AB rumble may be felt through the pilot seat. The
rumble does not impact engine operation and should
not be considered abnormal.
Inlet Thump GE 129
An inlet thump is an airflowrelated phenomenon
which may occur occasionally during PRI operation
below 25,000 feet and below 500 knots when thethrottle is retarded from MIL. The thump is the resultof airflow changes within the aircraft inlet.
A thump can be heard or can be felt through the
cockpit floor, ejection seat, and/or rudder pedals. The
intensity can vary greatly, ranging from a barely
perceptible sound or impulse to a louder and/or
harder occurrence which may be quite noticeable.
Thumps do not affect engine or aircraft operation.
SEC Operation GE 129
The engine can automatically transfer to SEC whena DEC failure occurs. It can manually be transferred
to SEC by placing the ENG CONT switch to SEC.
Transfer is indicated by the illumination of the SEC
caution light; rpm may also initially decrease (up to
10 percent rpm) and then recover to a level slightly
below that for PRI. When operating in SEC, thenozzle is closed.
Refer to ENGINE LIMITATIONS, Section V for
throttle restrictions while operating in SEC.
Movement of the throttle to MAX AB is permitted;
however, since the AB is inhibited, maximum
available thrust in SEC is attained at MIL. The
thrust level at MIL during SEC operation is 7095
percent of that available during PRI operation at thesame throttle position. Idle thrust in SEC is higher
than idle thrust in PRI because the exhaust nozzle isclosed. VSV reset is not active in SEC.

T.O. GR1F16CJ1AB

Light-Off Characteristics
ENGINE F110-GE-129

NOTES:
Regions 1 and 2

Unrestricted throttle movement. AB blowouts should not occur.

Region 1

Normal AB light-offs are expected.

Region 2

No lights or delayed lights accompanied by nozzle fluctuations indicating recyclingof AB initiation are possible.

70

60

REGION 2

ALTITUDE

1000 FEET

50

40

REGION 1
30

20

10

0
0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

2.0

2.2

MACH NUMBER

1F-16X-1-4007X

Figure 117.

157

T.O. GR1F16CJ1FIRE AND OVERHEAT DETECTIONSYSTEM
The fire and overheat detection system consists of two
separate parallel loop sensing systems, one for fire
and the other for overheat. The fire detection loops
are routed through the engine compartment. The
overheat detection loops are routed through theengine compartment, MLG wheel wells, ECS bay, andEPU bay. Activation of the overheat detection loops
occurs approximately 100_C below the activation
temperature of the fire detection loops. The firewarning signal causes the ENG FIRE warning lightto illuminate. The overheat signal causes the
OVERHEAT caution light to illuminate. When thetemperature of the element drops below the criticaltemperatures, the signal ceases, allowing the ENG
FIRE warning or the OVERHEAT caution light to go
off. The detection circuit is powered by emergency ac
bus No. 2 and battery bus No. 2.
FIRE & OHEAT DETECT TEST BUTTON C

DFRefer to figure 118. The FIRE & OHEAT DETECT
test button, located on the TEST switch panel, checks
continuity of both systems and illuminates the ENG
FIRE warning light and the OVERHEAT caution
light and provides a CSFDR special event data save
if depressed in flight.

FUEL SYSTEM
Refer to figure 119 for a simplified system diagramand figures 120 and 121 for system schematics. The
fuel system is divided into seven functional
categories. These are the fuel tank system, fueltransfer system, fuel tank vent and pressurization
system, engine fuel supply system, fuel quantity/fuellevel sensing system, fuel tank explosion suppression
system, and refueling/defueling system.
FUEL TANK SYSTEM
Refer to figure 122 for tank locations and capacities.
The aircraft has seven internal fuel tanks located in
the fuselage and wings that are integral to thestructure. There are provisions for carrying threeexternal tanks on the wings and the centerline
station PX III and for mounting two Conformal FuelTanks (CFT's) to the upper surface of the aircraft.
CFT fuel is considered internal fuel. Five of theinternal tanks are storage tanks: the left and right158TEST Switch Panel
C DF (Typical)

8
7
6

1
2
3
4

5
1. MAL & IND LTS Test Button
2. OXY QTY Indicator Test Switch
3. FLCS PWR Lights (Green)
4. FLCS PWR TEST Switch
5. DF STICK CONTROL Switch
6. EPU/GEN Test Switch
7. PROBE HEAT Switch
8. FIRE & OHEAT DETECT Test Button

GR1F-16CJ-1-0023X37

Figure 118.

wing tanks, two forward fuselage tanks (F1 and F2),and the aft fuselage tank (A1). The two internal
reservoir tanks (forward and aft) supply fuel directly
to the engine. D The F1 fuel tank is reduced in size
to allow room for the rear cockpit.
FUEL TRANSFER SYSTEM
Fuel is transferred by two independent methods. The
primary method provides a siphoning action throughstandpipes connecting the fuel tanks. Siphoning
action depends on the absence of air in the bays
receiving fuel. Air ejectors in each reservoir tank
automatically expel air. In case of failure of thesiphoning system, powered fuel pumps work
continually to pump fuel from the internal tanks to
the reservoirs. The powered transfer system also
scavenges tanks to minimize unusable fuel by using
electrically driven pumps and pumps powered bybleed fuel pressure from the engine manifold. Bothmethods operate simultaneously and independently
to transfer fuel through the system.

T.O. GR1F16CJ1

Fuel Tank System (Typical) PX IICENTERLINE
FUEL TANKAR RECEPTACLE
LEFT EXTERNAL
FUEL TANKF-1
GROUNDREFUELING
RECEPTACLE

RIGHT EXTERNAL
FUEL TANK

1

F-2

6

AFT RSVR

T

7
3

2

FWD
RSVR

T

4
RIGHT WING TANKLEFT WING TANKFUEL FLOWPROPORTIONER
HYD
A

5

LEGEND:

A-1

MAIN GEN CSD
FUEL FLOW
ADG
TRANSMITTER

HYD
B

MAIN
FUEL SHUTFUEL/OIL HEAT OFF VALVEEXCHANGER

BOOST PUMP PRESS
FUEL TRANSFER
AFT/LEFT FUEL TANK SYSTEM
FWD/RIGHT FUEL TANK SYSTEMTOENGINEELECTRICAL

EXTERNAL FUEL TANKFUEL TANK CONNECTING STANDPIPE

TRANSFER/REFUEL PRESSURIZATION

TRANSFER PUMP (FUEL PRESSURE DRIVEN)REFUEL/TRANSFER

TRANSFER PUMP (ELECTRICAL)DISCONNECT
T

TRANSFER SHUTOFF VALVEORFUEL FLOW DIRECTION

BOOST PUMP (ELECTRICAL)REFUEL SHUTOFF VALVESTANDPIPE OPENING AT TOP OF TANKSTANDPIPE OPENING AT BOTTOM OF TANKNOTE:This is a simplified C diagram. Not all fuel systemitems are shown (crossfeed, motive flow, etc). Refer
to FUEL SYSTEM SCHEMATIC for details.
GR1F-16CJ-1-0024A37

Figure 119. (Sheet 1)

159

T.O. GR1F16CJ1

Fuel Tank System (Typical) PX III
LEFT CFTRIGHT CFTCENTERLINE
FUEL TANKAR RECEPTACLE

LEFT EXTERNAL
FUEL TANKRIGHT EXTERNAL
FUEL TANKF-1

GROUNDREFUELING
RECEPTACLE

1

F-2

6

AFT RSVR

7
3

2

FWD
RSVR

4
T

RIGHT WING TANKLEFT WING TANKTFUEL FLOWPROPORTIONER
HYD
A

5

LEGEND:

A-1

MAIN GEN CSD
FUEL FLOW
ADG
TRANSMITTER

HYD
B

MAIN
FUEL SHUTFUEL/OIL HEAT OFF VALVEEXCHANGER

BOOST PUMP PRESS
FUEL TRANSFER
AFT/LEFT FUEL TANK SYSTEM
FWD/RIGHT FUEL TANK SYSTEMELECTRICAL

EXTERNAL FUEL TANKFUEL TANK CONNECTING STANDPIPE

TRANSFER/REFUEL PRESSURIZATION

TRANSFER PUMP (FUEL PRESSURE DRIVEN)REFUEL/TRANSFER

TRANSFER PUMP (ELECTRICAL)DISCONNECT
T

TRANSFER SHUTOFF VALVEORFUEL FLOW DIRECTION

BOOST PUMP (ELECTRICAL)REFUEL SHUTOFF VALVESTANDPIPE OPENING AT TOP OF TANKSTANDPIPE OPENING AT BOTTOM OF TANKNOTE:This is a simplified C diagram. Not all fuel systemitems are shown (crossfeed, motive flow, etc). Refer
to FUEL SYSTEM SCHEMATIC for details.

Figure 119. (Sheet 2)160TO
ENGINE

GR1F-16CJ-1-1024X37

T.O. GR1F16CJ1
The transfer system is divided into two separate tanksystems, the forward and the aft. The forward system
consists of the right external tank (if installed), PX III
right CFT (if installed), right internal wing tank, F1,
F2, and the forward reservoir. The aft system
consists of the left external tank (if installed), PX III
left CFT (if installed), left internal wing tank, A1,and the aft reservoir. If a centerline tank is installed,
it is considered to be part of both forward and aft
systems. The wing external tanks empty into therespective internal wing tanks. PX III The right CFT
transfers into the right internal wing tank, and theleft CFT transfers into the left internal wing tank.

The automatic forward fuel transfer system supple
ments the function of the FFP by preventing
undesirable aft CG. The system operates only when the
FUEL QTY SEL knob is in NORM and the total
forward fuselage fuel quantity indication is less than2800( D 1500)pounds. In the C , forward fueltransfer starts when the forward heavy fuel differential
drops below 300 pounds and stops when the forward
heavy fuel differential reaches 450 pounds. In the D ,forward fuel transfer starts when the aft heavy
differential exceeds 900 pounds and stops when the aftheavy fuel differential reaches 750 pounds. This systemdoes not correct a forward fuel imbalance since it onlytransfers fuel from aft to forward.

Fuel flows from the internal wing tanks to the fuselage
tanks and then to the forward and aft reservoirs. Fuelis pumped to the engine from the reservoirs. To
automatically maintain the CG, fuel is transferredthrough the forward and aft systems simultaneously.

For proper operation, the automatic forward fueltransfer system depends on a properly functioning
fuel quantity indicating system. Fuel is transferredthrough a solenoidoperated trim valve powered fromemergency dc bus No. 2. The automatic system isdeactivated if electrical power is lost through failure,by moving the FUEL QTY SEL knob out of NORM, orduring gravity feed conditions.

PX III The CFT fuel system gravity feeds fuel to theinternal wing tanks. Primary fuel transfer is bysiphon/air pressure transfer from the CFT to theinternal wing. CFT's will empty prior to the internal
wing tanks.

If external tanks are installed, air pressure transfers
fuel to the internal wing tanks. If the EXT FUEL
TRANS switch is in NORM, the sequence of fuel flow
is from the centerline tank to the internal wing tanks.
After the centerline tank empties, each external wingtank flows to its respective internal wing tank. The
external tank fuel transfer valve in each internal
wing tank shuts off fuel to prevent overfilling theinternal tanks. If one of these valves fails, a float
switch senses fuel and shuts off all external tank fueltransfer before fuel flows overboard. By placing the
EXT FUEL TRANS switch to WING FIRST, theexternal wing tanks empty before the centerline tank,and the float switch does not prevent fuel fromspilling overboard if a transfer valve fails.
PX III When using CFT's and external tanks,
selection of the EXT FUEL TRANS switch to CFTFIRST/NO FILL will deplete the CFT's prior to theexternal tanks. After the CFT's deplete, the transfer
sequence of the external tanks will be centerline first,
followed by the external wing tanks. When using
CFT's and external tanks, selection of the EXT FUEL
TRANS switch to NORM or WING FIRST will allow
air pressure transfer of fuel from the external tanks
into the internal wing tanks, keeping the CFT's andthe internal wing tanks full. The sequence of externaltank transfer will be as described in the previous
paragraph on external tank transfer. When externalfuel has transferred, CFT fuel will begin to transfer.

FUEL TANK VENT AND PRESSURIZATION SYSTEM
The fuel tank vent and pressurization system supplies
cooled pressurized air from the ECS to force fuel fromthe external tanks to the internal wing tanks and topower the air ejector pumps whenever the AIR
SOURCE knob is in NORM or DUMP. It also prevents
fuel in internal tanks from vaporizing at high altitude.
An external tank vent and pressurization valve
regulates pressure supplied to the external tanks.
If the combat schedule (reduced pressure) is activatedby the TANK INERTING switch, Halon, if available,
is mixed with air and the internal tank vent andpressurization valve controls the pressure.
If the AIR SOURCE knob is placed in OFF or RAM or
if the ECS is inoperative, tank pressurization is notavailable and external fuel cannot be transferred.
With multiple generator failures, fuel tank pressur
ization continues and external fuel still transfers.
ENGINE FUEL SUPPLY SYSTEM
Refer to figure 123 for fuel system controls andoperation. When the ENG FEED knob is in NORM,
boost pumps in the forward and aft reservoirs pumpthe fuel through the engine feedline to the fuel flow
proportioner (FFP). In the FFP, twin constantdisplacement pumps, powered by hydraulic system A,
supply equal amounts of fuel from each reservoir tomaintain CG. Two fuel lines with check valves can
bypass the FFP in case it fails so that fuel flow will not
be interrupted. After fuel flows through the FFP, asmall amount of cooling fuel is routed to the PW 229
161

T.O. GR1F16CJ1

Fuel System C (Typical) PX II

*

RIGHT
WING

*

T
7

**

**

4
3

FWD
RSVR

1

B

C
E

A

F-1

D
FLOAT
SWITCH

F-2

AFT
RSVR

2

TANK
INERTING
SWITCH

6
FROM ECSHALON RSVR

TRANSFERSWITCHT

FUEL
QUANTITY
INDICATORFUEL QTYSEL PANELEXTERNAL TANK(TYPICAL)

LEFT
WING
GR1F-16CJ-1-0026-1A37

Figure 120.(Sheet 1)162

T.O. GR1F16CJ1

ENG FEEDKNOBFUEL MASTER
SWITCHFUEL FLOW
INDICATOR

PUMP
DRIVE
PRESSURE
FROM
HYD
SYS A

LEGEND:
TRANSFER SHUTOFF VALVET
TRANSFER PUMP (ELECTRICAL)BOOST PUMP (ELECTRICAL)FUEL FLOW
TRANSMITTERREFUEL SHUTOFF VALVESOLENOID VALVEFUEL FLOWPROPORTIONERCROSSFEED VALVECAUTION LIGHT
FUEL/OILHOT

DEC

TRANSFER PUMP(FUEL PRESSURE DRIVEN)

TO JFSSUCTION FEED VALVEDISCONNECT

TO ENGINEGROUND COOLING RECEPTACLE
FUEL EJECTOR PUMP

FUEL/OIL HEATEXCHANGER

ELECTRIC MOTOR OPERATEDSHUTOFF VALVEORIFICECHECK VALVE

TRANSFER
TRIM
VALVE
AIR
REFUEL
SWITCHTEMPERATURE SENSOR
STANDPIPE OPENING AT TOP
OF TANK
STANDPIPE OPENING AT
BOTTOM OF TANKFUEL TANK CONNECTING
STANDPIPE

A-1

5

*
**
A
B
C
D
E

TO RIGHT WING EXT TANKTO CENTERLINE TANKHALON/AIR MIXING VALVEEXT TANK VENT & PRESS VALVEINT TANK VENT & PRESS VALVEREMOTE SENSING PRESSURE
RELIEF VALVEAR RECEPTACLE
ELECTRICALOVERBOARD VENTFUEL SUPPLYBOOST PUMP PRESSURE
PUMP DRIVE PRESSUREGROUND REFUELING
RECEPTACLETANK INERTING PRESSURE
FUEL TRANSFER
TRANSFER/REFUELPRESSURIZATIONAR STATUS
INDICATORTANK PRESSURIZATIONREFUEL/TRANSFERECS PRESSURIZATION SUPPLYFWDFUEL LOW
AFT
FUEL LOW

TRANSFER BLEEDCAUTION LIGHTS
GR1F-16CJ-1-0026-2X37

Figure 120.(Sheet 2)

163

T.O. GR1F16CJ1

Fuel System C (Typical) PX III

RIGHT
WING

***

*

*

T
7

**

**

***

4
3

FWD
RSVR

1

B

C
E

A

F-1

D
FLOAT
SWITCH

F-2

AFT
RSVR

2

CFT DRYPAN
TANK
INERTING
SWITCH

6
FROM ECSHALON RSVR

TRANSFERSWITCHT

FUEL
QUANTITY
INDICATORFUEL QTYSEL PANELEXTERNAL TANK(TYPICAL)

LEFT
WING

GR1F-16CJ-1-1026-1X37

Figure 120.(Sheet 3)164

T.O. GR1F16CJ1
ENG FEEDKNOBFUEL MASTER
SWITCHFUEL FLOW
INDICATOR

PUMP
DRIVE
PRESSURE
FROM
HYD
SYS A

LEGEND:
TRANSFER SHUTOFF VALVET
TRANSFER PUMP (ELECTRICAL)BOOST PUMP (ELECTRICAL)FUEL FLOW
TRANSMITTERREFUEL SHUTOFF VALVESOLENOID VALVEFUEL FLOWPROPORTIONERCROSSFEED VALVECAUTION LIGHT
FUEL/OILHOT

DEC

TRANSFER PUMP(FUEL PRESSURE DRIVEN)

TO JFSSUCTION FEED VALVEDISCONNECT

TO ENGINEGROUND COOLING RECEPTACLE
FUEL EJECTOR PUMP

FUEL/OIL HEATEXCHANGER

ELECTRIC MOTOR OPERATEDSHUTOFF VALVEORIFICECHECK VALVE

TRANSFER
TRIM
VALVE

A-1

AIR
REFUEL
SWITCH

*
**
** *
A
B
C
D

5

E
F
AR STATUS
INDICATORTEMPERATURE SENSOR
STANDPIPE OPENING AT TOP
OF TANK
STANDPIPE OPENING AT
BOTTOM OF TANKFUEL TANK CONNECTING
STANDPIPE
TO RIGHT WING EXT TANKTO CENTERLINE TANKTO RIGHT CFT
HALON/AIR MIXING VALVEEXT TANK VENT & PRESS VALVEINT TANK VENT & PRESS VALVEREMOTE SENSING PRESSURE
RELIEF VALVEAR RECEPTACLEREMOTE SENSING PRESSURE
RELIEF VALVE (CFT)ELECTRICALOVERBOARD VENTFUEL SUPPLYGROUND REFUELING
RECEPTACLEBOOST PUMP PRESSURE
PUMP DRIVE PRESSURE
TO VENT TANKFWDFUEL LOWFTANK INERTING PRESSURE
FUEL TRANSFER
TRANSFER/REFUELPRESSURIZATIONAFTFUEL LOWTANK PRESSURIZATION
8CAUTION LIGHTSTO WINGCONFORMAL FUEL TANK(TYPICAL)REFUEL/TRANSFERECS PRESSURIZATION SUPPLYTRANSFER BLEED
GR1F-16CJ-1-1026-2X37

Figure 120.(Sheet 4)

165

T.O. GR1F16CJ1

Fuel System D (Typical) PX II

*

RIGHT
WING

*

T
7

**

**
4
FWD
RSVR

B

C
E

A

F-1

D
FLOAT

F-2

SWITCH

3

AFT
RSVR

2

1

TANK
INERTING
SWITCH

6
FROM ECSHALON RSVR

TRANSFERSWITCHFUEL QTYTDF SEL PANEL

EXTERNAL TANK(TYPICAL)
FWD & AFT FUELQUANTITY INDICATORS

GR1F-16CJ-1-0027-1A37

Figure 121.(Sheet 1)166

LEFT
WING

T.O. GR1F16CJ1
FWD FUELMASTER SWITCHAFT FUELMASTER SWITCH

FWD & AFT FUELFLOW INDICATORLEGEND:T TRANSFER SHUTOFF VALVE
TRANSFER PUMP (ELECTRICAL)BOOST PUMP (ELECTRICAL)ENG FEED
KNOB
PUMP
DRIVE
PRESSURE
FROM HYDSYS AREFUEL SHUTOFF VALVESOLENOID VALVEFUEL FLOWPROPORTIONER

FWD & AFT
CAUTION LIGHT

CROSSFEED VALVE

FUEL/OILHOT

DEC

TRANSFER PUMP(FUEL PRESSURE DRIVEN)
SUCTION FEED VALVEDISCONNECTGROUND COOLING RECEPTACLETO ENGINE
TO JFSFUEL EJECTOR PUMPFUEL FLOW TRANSMITTER
FUEL/OIL HEAT EXCHANGERELECTRIC MOTOR OPERATEDSHUTOFF VALVEORIFICECHECK VALVE

TRANSFER
TRIM
VALVE
AIR
REFUEL
SWITCH

A-1
*
**
A
B
C
D

5

E

TEMPERATURE SENSOR
STANDPIPE OPENING AT TOP
OF TANK
STANDPIPE OPENING AT
BOTTOM OF TANKFUEL TANK CONNECTING
STANDPIPE
TO RIGHT WING EXT TANKTO CENTERLINE TANKHALON/AIR MIXING VALVEEXT TANK VENT & PRESS VALVEINT TANK VENT & PRESS VALVEREMOTE SENSING PRESSURE
RELIEF VALVEAR RECEPTACLE
ELECTRICALOVERBOARD VENTFUEL SUPPLYBOOST PUMP PRESSURE
PUMP DRIVE PRESSUREGROUND REFUELING
RECEPTACLE

FWD ARSTATUS
INDICATORTANK INERTING PRESSURE
FUEL TRANSFER
TRANSFER/REFUELPRESSURIZATION

AFT ARSTATUS
INDICATORTANK PRESSURIZATIONREFUEL/TRANSFERECS PRESSURIZATION SUPPLYTRANSFER BLEED

FWD
FUEL LOW
AFT
FUEL LOWFWD CAUTION
LIGHTS

AFT CAUTION
LIGHTS

FWD
FUEL LOW
AFT
FUEL LOW

GR1F-16CJ-1-0027-2X37

Figure 121.(Sheet 2)

167

T.O. GR1F16CJ1

Fuel System D (Typical) PX III

RIGHT
WING

***

*

*

T
7

**

**
***

4
FWD
RSVR

B

C
E

A

F-1

D
FLOAT

F-2

SWITCH

3

AFT
RSVR

2

1
CFT DRYPAN
TANK
INERTING
SWITCH

6
FROM ECSHALON RSVR

TRANSFERSWITCHFUEL QTYTDF SEL PANEL

EXTERNAL TANK(TYPICAL)
FWD & AFT FUELQUANTITY INDICATORS

GR1F-16CJ-1-1027-1X37

Figure 121.(Sheet 3)168

LEFT
WING

T.O. GR1F16CJ1
FWD FUELMASTER SWITCHAFT FUELMASTER SWITCH

FWD & AFT FUELFLOW INDICATORLEGEND:T TRANSFER SHUTOFF VALVE
TRANSFER PUMP (ELECTRICAL)BOOST PUMP (ELECTRICAL)ENG FEED
KNOB
PUMP
DRIVE
PRESSURE
FROM HYDSYS AREFUEL SHUTOFF VALVESOLENOID VALVEFUEL FLOWPROPORTIONER

FWD & AFT
CAUTION LIGHT

CROSSFEED VALVE

FUEL/OILHOT

DEC

TRANSFER PUMP(FUEL PRESSURE DRIVEN)
SUCTION FEED VALVEDISCONNECTGROUND COOLING RECEPTACLETO ENGINE

FUEL EJECTOR PUMPFUEL FLOW TRANSMITTERTO JFS

FUEL/OIL HEAT EXCHANGERELECTRIC MOTOR OPERATEDSHUTOFF VALVEORIFICECHECK VALVE

TRANSFER
TRIM
VALVE

A-1

AIR
REFUEL
SWITCHFWD ARSTATUS
INDICATOR

*
**
** *
A
B
C
D

5

E
F

AFT ARSTATUS
INDICATORTEMPERATURE SENSOR
STANDPIPE OPENING AT TOP
OF TANK
STANDPIPE OPENING AT
BOTTOM OF TANKFUEL TANK CONNECTING
STANDPIPE
TO RIGHT WING EXT TANKTO CENTERLINE TANKTO RIGHT CFT
HALON/AIR MIXING VALVEEXT TANK VENT & PRESS VALVEINT TANK VENT & PRESS VALVEREMOTE SENSING PRESSURE
RELIEF VALVEAR RECEPTACLEREMOTE SENSING PRESSURE
RELIEF VALVE (CFT)ELECTRICALOVERBOARD VENTFUEL SUPPLYGROUND REFUELING
RECEPTACLEBOOST PUMP PRESSURE
PUMP DRIVE PRESSURE
TO VENT TANKFWDFUEL LOW

AFT CAUTION
LIGHTS FUELAFTLOWFWDFUEL LOW
AFT
FUEL LOWTANK INERTING PRESSURE
FUEL TRANSFERF

TRANSFER/REFUELPRESSURIZATIONTANK PRESSURIZATIONFWD CAUTION
LIGHTS

8
TO WINGREFUEL/TRANSFERECS PRESSURIZATION SUPPLYCONFORMAL FUEL TANK(TYPICAL)

TRANSFER BLEED
GR1F-16CJ-1-1027-2X37

Figure 121.(Sheet 4)

169

T.O. GR1F16CJ1

Fuel Quantity Indication and Tank
Arrangement (Typical) PX II 2 4
5

6

7

10
1
3
C

D

FUEL QUANTITY INDICATIONS (JP-8) WITHNORM SELECTED AND FULL INTERNALTANKS WITH TWO FULL 370-GALLON
EXTERNAL WING TANKS.

98

GR1F-16CJ-1-0025A37DC
TANK LOCATION

1. LEFT INTERNAL WING
2. RIGHT INTERNAL WING

FUELQTYSELKNOB SET POINTER
TINGS

FUEL
QTY (LB) JP4FUEL
QTY (LB) JP5/8

FUEL
QTY (LB) JP4FUEL
QTY (LB) JP5/8

INT WING
INT WING

AL
FR

525$100
525$100

550$100
550$100

525$100
525$100

550$100
550$100

3. F1 FUSELAGE
4. F2 FUSELAGE
5. FWD RESERVOIR

NORM

FR

3100$100

3250$100

1800$100

1890$100

6. AFT RESERVOIR
7. A1 FUSELAGE

NORM

AL

2675$100

2810$100

2675$100

2810$100

5. FWD RESERVOIR
6. AFT RESERVOIR

RSVR
RSVR

FR
AL

460$30
460$30

480$30
480$30

460$30
460$30

480$30
480$308. CENTERLINE
9. LEFTEXTERNALWING
370/600

EXT CTR
EXT WING

FR
AL

1800$100
2300/3750$100

1890$100
2420/3925$100

1800$100
2300/3750$100

1890$100
2420/3925$100

10. RIGHTEXTERNALWING
370/600

EXT WING

FR

2300/3750$100

2420/3925$100

2300/3750$100

2420/3925$100

TOTAL INTERNAL FUEL

6825$300

7160$300

5575$300

5800$300

TOTAL EXTERNAL FUEL (370/600)

6400/9300$300

6730/9740$300

6400/9300$300

6730/9740$300

NOTES:
1. Tolerances are due to indication errors with the variations in density resulting from temperatures, additives, etc.
2. The quantity of wing fuel varies depending upon aircraft attitude during refueling.
3. Usable fuel and indicated fuel quantities are approximately equal.
4. Indications are approximate and shall not be used for computing weight and balance data. Refer to T.O. GR1F16CJ11 for detailed
information.

Figure 122. (Sheet 1)170Change 1

T.O. GR1F16CJ1

Fuel Quantity Indication and Tank CFT
Arrangement (Typical) PX III 2
7
4

5

6

10
CFT 1
3
C

D

FUEL QUANTITY INDICATIONS (JP-8) WITHNORM SELECTED AND FULL INTERNALTANKS WITH TWO FULL 370-GALLON
EXTERNAL WING TANKS.

98

GR1F-16CJ-1-1025X37DC
TANK LOCATION

FUELQTYSELKNOB SET POINTER
TINGS

FUEL
QTY (LB) JP4FUEL
QTY (LB) JP5/8

FUEL
QTY (LB) JP4FUEL
QTY (LB) JP5/8

1. LEFT INTERNAL
WING/WITH CFT
2. RIGHT INTERNAL
WING/WITH CFT

INT WING& CFT
INT WING& CFT

AL

525/1965$100

550/2080$100

525/1965$100

550/2080$100

FR

525/1965$100

550/2080$100

525/1965$100

550/2080$100

3. F1 FUSELAGE
4. F2 FUSELAGE
5. FWD RESERVOIR

NORM

FR

3100$100

3250$100

1800$100

1890$100

6. AFT RESERVOIR
7. A1 FUSELAGE

NORM

AL

2675$100

2810$100

2675$100

2810$100

5. FWD RESERVOIR
6. AFT RESERVOIR

RSVR
RSVR

FR
AL

460$30
460$30

480$30
480$30

460$30
460$30

480$30
480$30EXT CTR
EXT WING

FR
AL

1800$100
2300/3750$100

1890$100
2420/3925$100

1800$100
2300/3750$100

1890$100
2420/3925$100

EXT WING

FR

2300/3750$100

2420/3925$100

2300/3750$100

2420/3925$100

TOTAL INTERNAL FUEL (Without CFT's/With CFT's)

6825/9705$300

7160/10220$300

5525/8405$300

5800/8860$300

TOTAL EXTERNAL FUEL (370/600)

6400/9300$300

6730/9740$300

6400/9300$300

6730/9740$300

8. CENTERLINE
9. LEFTEXTERNALWING
370/600
10. RIGHTEXTERNALWING
370/600

NOTES:
1. Tolerances are due to indication errors with the variations in density resulting from temperatures, additives, etc.
2. The quantity of wing fuel varies depending upon aircraft attitude during refueling.
3. Usable fuel and indicated fuel quantities are approximately equal.
4. Indications are approximate and shall not be used for computing weight and balance data. Refer to T.O. GR1F16CJ11 for detailed
information.
5. Actual reservoir quantities are 23 pounds less than indicated because of provisions for CFT's.

Figure 122. (Sheet 2)

171

T.O. GR1F16CJ1

Fuel Control Panel (Typical)
C

1

2

DF

DR

4

3

1
1. FUEL MASTER Switch
2. TANK INERTING Switch
3. ENG FEED Knob
4. AIR REFUEL Switch
GR1F-16CJ-1-0028X37

1.

2.

CONTROLPOSITIONFUNCTION

FUEL MASTER Switch(lever lock)

MASTER (guarded)

Opens main fuel shutoff valve which thenopens the engine electronic control cooling fuelshutoff valve

OFF

Closes main fuel shutoff valve which thencloses the engine electronic control cooling fuelshutoff valve

TANK INERTING

Reduces internal tank pressurization. If
Halon is available, allows PX II 20, PX III
30 seconds of initial Halon flow; thereafter,allows a small metered flow of Halon to
the F1, A1, internalwingtanks,and PX III
CFTs

OFF

Stops Halon flow. Returns internal tankpressurization to normal schedule

TANK INERTING Switch(lever lock to OFF)

Figure 123.(Sheet 1)172

T.O. GR1F16CJ1

Fuel Control Panel (Typical)
CONTROL
3.

4.

ENG FEED Knob

AIR REFUEL Switch

POSITIONFUNCTION

OFF

Deenergizes all electricdriven pumps. Engine
supplied by FFP

NORM

Energizes all pumps. CG maintained auto
matically

AFT

Energizes pumps in aft tanks and opens cross
feed valve. Fuel is transferred from aft tanks
to the engine and forward tanks. CG moves
forward

FWD

Energizes pumps in forward tanks and opens
crossfeed valve. Fuel is transferred from for
ward tanks to the engine and aft tanks. CG
moves aft

OPEN

Opens slipway door. Places FLCS in takeoff
and landing gains when airspeed is below 400
knots
Enables slipway light. Turns on AR floodlightand vertical tailmounted fuselage floodlight
Reduces internal tank pressurization,
depressurizes external tanks, and allows therefuel valve in each reservoir to open when acenterline tank is installed and refuel pres
sure is applied

CLOSE

Reverses the OPEN actions

Figure 123.(Sheet 2)

173

T.O. GR1F16CJ1
DEEC, GE 129 DEC and then returned to thereservoirs. The remainder of the fuel passes througha fuel/oil heat exchanger to cool hydraulic systems A
and B, the main generator CSD, and the ADG. Thenfuel flows through an electric main fuel shutoff valve
which has a full travel time of 24 seconds and iscontrolled by the FUEL MASTER switch. (The JFS
receives fuel at all times regardless of the FUELMASTER switch position.) After passing through themain fuel shutoff valve, fuel passes through the fuelflow transmitter (which operates the FUEL FLOW
indicator) to the engine.
FUEL QUANTITY INDICATING SYSTEM
Refer to figure 124. The fuel quantity indicating
system displays the amount and location of fuelremaining. The totalizer shows all fuel in theinternal and external tanks in pounds. The AL and
FR pointers show the fuel quantity in the tanks as
selected by the FUEL QTY SEL knob. Erroneous fuelindications may occur during or immediately after
maneuvering flight. The selected tanks should
normally be the fuselage tanks (FUEL QTY SEL
knob in NORM). The difference between the forward
and aft tanks should remain essentially constantsince the FFP maintains an equal flow of fuel. PX III
With the FUEL QTY SEL knob in the INT WING &CFT position, the pointers show the combined total
fuel in both tanks.
Normally, the forward tank fuel quantity is 0600
pounds greater than the aft tank quantity. D
Normally, the aft tank fuel quantity is 7001350
pounds greater than the forward fuel quantity. If
these values are exceeded in either direction, a red
portion of the AL pointer becomes visible. Fueldistribution can be changed by rotating the ENG
FEED knob to the FWD or AFT position until theimbalance is corrected.
C

RESERVOIR FUEL LEVEL SENSING SYSTEM
Fuel level sensors in the reservoir tanks are used toturn on/off the air ejectors and the fuel low caution
lights. When a reservoir tank is not full, the air
ejector in that tank is operating. The reservoir tanksensors, associated sensor circuitry, and fuel levelsensing unit operate independently of the fuelquantity indicating system.
Fuel Low Caution Lights
The fuel low caution lights, located on the caution
light panel, indicate either a low fuel quantity in thereservoir tanks or a reservoir fuel level sensing
system malfunction. The caution lights function
174

independently of the fuel quantity indicating system.
The FWD FUEL LOW caution light illuminates whenfuel quantity in the forward reservoir drops below 400( D 250) pounds. The AFT FUEL LOW caution lightilluminates when aft reservoir fuel quantity drops
below 250 ( D 400) pounds. The caution lights are
powered by emergency dc bus No. 1.
HUD FUEL LOW/BINGO INDICATION
In addition to the fuel low caution lights, a fuel low
condition may be indicated by the word FUEL in the
HUD in conjunction with the home mode of the FCC
or the previously entered bingo fuel value.
Bingo fuel warning is based either on fuselage fuelwith the FUEL QTY SEL knob in NORM or on total
fuel with the FUEL QTY SEL knob out of NORM.
With the FUEL QTY SEL knob in NORM, the bingo
computation is based on the lesser of fuselage fuelweight or total fuel weight. That is, with the FUELQTY SEL knob in NORM, bingo fuel warning will be
triggered when either fuselage fuel or total fueldecreases below the bingo fuel value.
With the FUEL QTY SEL knob out of NORM, thewarning will only be triggered when total fueldecreases below the bingo value. With trapped
external fuel, this could lead to fuel starvation before
the bingo warning is triggered.
The VMS provides a BINGOBINGO message in theheadset when the bingo fuel warning is activatedwith weight off wheels.
For a more detailed description of the home mode andthe bingo fuel option, refer to T.O. GR1F16CJ3411.
HUD TRP FUEL WARNING
A trapped external fuel condition is indicated byflashing TRP FUEL and FUEL in the HUD. Fiveconditions must be met for a TRP FUEL warning tooccur.
Conditions are:
D FUEL QTY SEL knob is in NORM.
D Aerial refueling has not occurred within previous

3090 seconds.

D Fuselage fuel has been at least 500 pounds lessthan fuselage capacity for 30 seconds.

D Total fuel has been at least 500 pounds greater thanfuselage fuel for 30 seconds.

D Fuel flow has been less than 18,000 pph for 30

seconds.

T.O. GR1F16CJ1

Fuel Quantity Indicator and Select Panel(Typical)
DF

C

DF

C

DR
1

2

3
PX II

CONTROL/INDICATOR
1.

2.

FUEL Quantity
Indicator

FUEL QTY SELKnobPX III

1. FUEL Quantity Indicator
2. FUEL QTY SEL Knob
3. EXT FUEL TRANS Switch

POSITION

GR1F-16CJ-1-0029A37

FUNCTION

AL and FR
pointers

Display fuel quantities as determined by the FUEL QTY SEL
knob

Totalizer

Displays total fuel in all fuel tanks (fuselage + wing + external).
The totalizer and the fuel value displayed on the DED BIN
GO page should agree within 100 pounds of each otherRed portion of
AL pointershowing

Indicates fuel imbalance between forward and aft fuselage
tanks

TEST

AL/FR pointers drive to 2000 ($100) pounds
Totalizer drives to 6000 ($100) pounds
Both fuel low caution lights illuminate
PX III (with CFT's installed) CFT pumps operate and continue
to operate for 30 seconds after switch is taken out of TEST
position; CFT pump indicator lights illuminate (ground test
panel)

Figure124. (Sheet 1)

175

T.O. GR1F16CJ1

Fuel Quantity Indicator and Select Panel(Typical)
CONTROL/INDICATOR
2.

FUEL QTY SELKnob (Cont)POSITIONNORM

FUNCTION
AL pointer displays sum of fuel in the aft (left) reservoir and A1
fuselage tanks
FR pointer displays sum of fuel in the forward (right) reservoir
and F1, F2 fuselage tanks
Enables automatic forward fuel transfer system, trapped fuelwarning, and bingo fuel computation based on fuselage fuel

RSVR
PX II

AL/FR pointers display fuel in aft/forward reservoir tanks
AL/FR pointers display fuel in left/right internal wing tanks

INT WING
PX III INTWING & CFT

AL pointer displays sum of fuel in the left CFT and leftinternal wing tank
FR pointer displays sum of fuel in the right CFT and rightinternal wing tank

EXT WING

AL/FR pointers display fuel in left/right external wing tanks

EXT CTRAL pointer drops to zero
FR pointer displays fuel in centerline tank3.

EXT FUEL
TRANS Switch

PX III CFTFIRST/NO FILL

On transfer, CFT's will transfer prior to external tanks.
On refuel (ground or air), CFT's will not fill

NORM

Centerline tank transfers first and then external wing tanks

WING FIRST

External wing tanks transfer first and then centerline tank

Figure 124. (Sheet 2)

176

T.O. GR1F16CJ1
A false TRP FUEL warning may occur after thefollowing:
D A fuel leak which exceeds the transfer rate of theexternal tank(s).

D Prolonged AB use if fuel flow to the engine exceeds

the transfer rate from the external tank(s).

D Receiving a partial fuel load during air refueling

Because of limited Halon supply, the system shouldbe activated after the external tanks have emptied,but before half of the internal fuel is depleted. PX III
Inerting may be selected at any CFT fuel state. Since
the PX II 20, PX III 30 seconds of initial inerting
occurs each time the TANK INERTING switch isplaced to TANK INERTING, do not cycle the switch.
The fuel tank explosion suppression system does notprotect the external fuel tanks.

with an external tank(s).

The TRP FUEL warning clears automatically after
the condition is corrected; the FUEL mnemonic may
be manually reset by placing the WARN RESET
switch to WARN RESET.
FUEL/OIL HOT CAUTION LIGHT
The FUEL/OIL HOT caution light, located on thecaution light panel, illuminates when the temperature of fuel to the engine becomes excessive. GE 129
The caution light also comes on as a function of hot oil.
Refer to ENGINE GE 129 , this section.
FUEL TANK EXPLOSION SUPPRESSION SYSTEM
The fuel tank explosion suppression system places
the fuel tank vent and pressurization system on a
reduced pressure schedule and inerts the fuel vapors
inside the tanks (if serviced with Halon). The system,
intended for use only in combat or during
emergencies, is controlled by the TANK INERTING
switch on the fuel control panel. The system uses
Halon as an inerting agent which prevents
combustion when mixed with air. For the agent
specification and reservoir location, refer to SERVICING DIAGRAM, this section. The Halon
reservoir has a heater, controlled by a thermostatic
switch, which assures sufficient operating pressure.
The RMLG WOW switch prevents operation of theheater while the aircraft is on the ground.
When the TANK INERTING switch is placed to
TANK INERTING, the fuselage tanks, internal wingtanks, PX III and CFT's (if present) are placed on a
reduced pressure schedule and a valve at the Halon
reservoir is opened. At each activation of the TANKINERTING switch, Halon (if available) is released
into the F1, A1, internal wing tanks, PX III and
CFT's (if present) for PX II 20, PX III 30 seconds for
initial inerting. Thereafter, a continuous meteredflow of Halon is mixed with the pressurization air tomaintain the inert condition. The metered flow
continues until the system is turned off or until the
MAIN PWR switch is positioned to OFF.

REFUELING SYSTEM
Ground Refueling
All external and internal fuel tanks can be pressure
filled from a singlepoint ground refueling receptacle
located on the lower left side of the fuselage justforward of the wing trailing edge. Electrical power is
not required to refuel the aircraft unless fuel quantity
is to be monitored or PX III CFT's are mounted and it
is desired that they not be filled, in which case power
must be applied and the EXT FUEL TRANS switch
must be placed in the CFT FIRST/NO FILL position.
Terminating refueling with partially filled tanks
could result in fuel imbalance. When a partial fuelload is required, fuel distribution should be correctedprior to flight by selective operation of the fueltransfer pumps controlled by the ENG FEED knob.
Air Refueling (AR) System
The AR system consists of a hydraulically actuatedreceptacle and slipway door, a signal amplifier, andthe associated controls and indicators. Hydraulicsystem B provides pressure for operation of the door
and latch mechanism. The receptacle is located on the
top fuselage centerline aft of the canopy. When theslipway door is opened, a mechanical linkage retracts
the aft end of the slipway door into the fuselage,
forming a slipway into the receptacle.
When the AIR REFUEL switch is placed to OPEN,
the external tanks are depressurized, external fueldoes not transfer, and the FLCS is placed in takeoff
and landing gains if airspeed is below 400 knots.
When closed, the slipway door is flush with thefuselage skin. The AR receptacle is equipped withfour lights, two located on each side. An AR floodlight
is located on the top fuselage centerline immediately
aft of the canopy. A light on the upper leading edge ofthe vertical tail floods the AR receptacle area and theupper fuselage.

177

T.O. GR1F16CJ1During AR operations, the AR boom enters thereceptacle and is automatically latched in place by ahydraulic actuating mechanism. The HOT MIC
switch allows intercom communications with com
patible tankers through the AR boom. When the lastrefuel shutoff valve closes, a pressure switch
automatically provides a signal to unlatch the boom
from the receptacle. A disconnect signal can be
manually initiated at anytime during AR by thereceiver or by the tanker boom operator.
D Disconnect from the boom may occur before alltanks are full if the external fuel tank configurationconsists of only a centerline fuel tank. Such a
disconnect typically occurs when refueling with an
initial internal fuel load of 4000 pounds or more andthe centerline tank empty. At disconnect, the aircraft
total fuel may be up to 1600 pounds less than full,with many occurrences resulting in approximately
1000 pounds less than full.
PX III When CFT's are present, the option to fill or
prevent filling of the CFT's is available. By placing
the EXT FUEL TRANS switch in CFT FIRST/NO
FILL, the CFT's will not fill. If the EXT FUEL TRANS
switch is in NORM or WING FIRST, the CFT's will
fill.

Fuel venting from under the left wing can occur
during AR, particularly when the aircraft isconfigured with external fuel tank(s). Terminating
the AR operation in a partially filled condition could
result in fuel imbalance. When a partial fuel load isrequired, fuel distribution should be monitored andcorrected as required by use of the ENG FEED knob.
NWS A/R DISC MSL STEP BUTTONThe NWS A/R DISC MSL STEP button is located on theoutboard side of the stick. The A/R DISC function of theswitch is activated when the aircraft is airborne and the
AIR REFUEL switch is positioned to OPEN. The
button provides a means of manually disconnecting the
AR boom. Depressing the switch causes the boom
latching mechanism to unlatch and release the boom.
AIR REFUELING (AR) STATUS INDICATOR
The AR status indicator, located to the right of the
HUD, contains three lights.

178

Functions are:
D RDY - Illuminates blue when the AR slipway door

is open and the system is ready.

D AR/NWS - Illuminates green when the boom islatched in place.

D DISC- Illuminates amber when a disconnect
occurs. After the disconnect, the system automati
cally recycles to ready, and the RDY lightilluminates after a 3second delay.

A lever for dimming the three lights is located on theright side of the unit.

ENVIRONMENTAL CONTROL SYSTEM(ECS)Refer to figure 125. The ECS combines aircondition
ing and pressurization functions to provide temperaturecontrolled, pressureregulated air for heating,cooling, ventilating, canopy defogging, cockpit
pressurization, canopy sealing, gsuit pressurization,
fuel tank pressurization, electronic equipment
cooling, pressure breathing for g (PBG), and PX IIIOBOGS. Most of these functions are lost when the
AIR SOURCE knob is placed to OFF or RAM. Refer
to AIR SOURCE KNOB, this section.
Above 35,000 feet MSL, automatic changes inoperation of the ECS may appear as short duration(approximately 15 seconds) losses of cockpit airflow
alternating with normal cockpit airflow levels. These
airflow changes are normal and are designed toprevent total ECS shutdowns.
ELECTRICAL FAILURES
When the ECS loses electrical power, cabintemperature control and water separator antiicing
become inoperative. The cockpit receives cold air only.
The water separator coalescer sock freezes ataltitudes where there is enough moisture in the air(below 30,000 feet) and the builtin bypass valve
opens allowing air to flow to the cockpit and avionics.
There is no interruption of bleed air flow due to a lossof electrical power. All pressurization functions (servo
air, canopy seal, antig, PBG, and fuel tankpressurization) operate normally since the bleed air
function continues to operate. PX III OBOGS also
continues to receive ECS air; however, a problem thatcauses loss of power to the ECS may also cause a lossof power to OBOGS.

T.O. GR1F16CJ1

Environmental Control System (Typical)
CAUTION LIGHT
CABIN
PRESS
COCKPITMECHANICALLY OPERATEDFLAPPER VALVERAVIONICEQUIPCAUTION LIGHT
EQUIP
HOT

C

ANTI-G VALVEAND CANOPYSEALM

SERVO AIR

E

C

M

C

M

M

R

M

FUEL TANKPRESSURIZATIONPX III OBOGS

T

C

R

EPU
ENGINECOMPRESSORBLEED AIRWS

LOW
PRESSURE

AIR OVERTEMPRELAYGROUNDCOOLING
RECEPTACLE

R

NACELLEEJECTORHIGHPRESSURE

R

R

E

TOTAL
TEMP
FROM
TT PROBE

Hx

R
GROUNDSERVICING
CONNECTIONE

RAM AIR

Hx
Hx

RAM AIRMECHANICALELECTRICALLEGEND:SOLENOID VALVECHECK VALVEM

MODULATING VALVE

TEMP/PRESS SENSOR

RPRESSURE REGULATORBLEED AIR
WARM AIR
COOL AIRRAM AIRCSENSOR/CONTROLLERE

ENVIRONMENTAL
DIGITAL CONTROL SET

WS WATER SEPARATOR
T/C TURBINE/COMPRESSOR
AIR-TO-AIR
Hx HEATEXCHANGER
GR1F-16CJ-1-0030A37

Figure 125.

179

T.O. GR1F16CJ1AIRCONDITIONING
Engine bleed air is directed through a turbine
compressor and airtoair heat exchangers where it iscooled by ram air. The conditioned air is then used forthe functions shown in figure 125.
A cockpit temperature controller receives signals
from temperature sensors and from a manuallyoperated control panel to automatically control the
cockpit temperature. Conditioned air enters the
cockpit on both sides, the top rear of the seat, throughthe angle vent on the instrument panel, and throughthe canopy defogger. In the event of an ECS
malfunction, emergency ram air operation can beselected for ventilation and cooling.

A ground cooling cart can be connected to the ground
cooling receptacle on the lower left side of the fuselage
just above the nosewheel area to provide cooling air
to the cockpit and avionic equipment.
PRESSURIZATION
Air pressure is provided by the pressurization systemfor control/operation of some of the ECS, canopy seal,gsuit, PBG, fuel tanks, radar and PX III OBOGS.
Pressure in the cockpit is controlled automaticallyaccording to the schedule shown in figure 126. A
cockpit pressure safety valve relieves pressure
anytime the cockpit pressure exceeds ambientpressure by 5.4 psi.
The canopy seal is inflated/deflated with themechanical locking/unlocking of the canopy.

Cockpit Pressure Schedule

COCKPIT ALTITUDE

1000 FEET

60

50

40

30

20

10

SL
SL

5

10

15

20

25

30

AIRCRAFT ALTITUDE

Figure 126.

180

35

40

1000 FEET

45

50

55

60

1F-16X-1-0001X

T.O. GR1F16CJ1AIR SOURCE Knob C

DF

Refer to figure 127. The AIR SOURCE knob islocated on the ECS panel.
Functions are:
D OFF

- Engine bleed air valves close. All
airconditioning, cooling, and pressurizing func
tions shut off, including gsuit, PBG, canopy seal,
fuel tank pressurization, and PX III OBOGS.

Environmental Control System Control
Panel C DF (Typical)

D NORM - Airconditioning system set for automatic

temperature and pressure regulation.

D DUMP - Cockpit pressure dump valve opens toatmospheric pressure. Cockpit pressure altitude
increases if DUMP is selected above approximately
8000 feet MSL. Conditioned air ventilates cockpit
and performs all other system functions.

D RAM - Engine bleed air valves close and the cockpit

pressure dump valve opens to atmosphericpressure. Cockpit pressure altitude increases if
RAM is selected above approximately 8000 feet
MSL. All airconditioning, cooling, and pressurizing
functions shut off, including gsuit, PBG, canopy
seal, fuel tank pressurization, and PX III OBOGS.
The ram air valve opens to admit ram air toventilate the cockpit and avionic equipment.

TEMP Knob C

1

2

1. TEMP Knob
2. AIR SOURCE Knob
GR1F-16CJ-1-0032X37

Figure 127.

DF

Refer to figure 127. The TEMP knob is located on the
ECS panel, only controls cockpit temperature.
Functions are:
D AUTO-Cockpit temperature is automatically main

tained (50_80_F) relative to the setting of the knob.

D MAN - The temperature control drives the air

modulating valve to a set position. Cockpit
temperature varies according to throttle setting,
OAT, and cockpit heat load. If WARM is selected, thecockpit supply air temperature may exceed the
maximum allowable limit of approximately 177_F.
This causes the warm air valve to cycle on and off.
This is a normal occurrence and can be stopped by
selecting a cooler setting.

D TEMP OFF - Hot air mixing is shut off. Only air at

approximately 35_F is delivered to cockpit.

Under extreme temperature conditions, system
performance on the ground can be improved byadvancing the throttle 13 percent above idle rpm.
Ground operation with the radar in OFF improves
cockpit cooling and ground operation with the radarin STBY improves cockpit heating.

DEFOG Lever C

DF

Refer to figure 128. The DEFOG lever, located on thefar aft portion of the left console, mechanicallycontrols a flapper valve in the cockpit air supply line.
Functions are:
D MIN - Minimum airflow toward the canopy forward

area and air vent in center pedestal; maximum
airflow to outlets behind seat.

D MAX - Most of the cockpit air supply is diverted tothe canopy forward area for defogging and to the air
vent in the center pedestal. A partial opening of thecenter pedestal air outlet allows a more balanced
defogging of the right and left sides of the canopy.
When placed in the full forward (MAX) defog
position and with the TEMP knob in AUTO, thelever activates a switch which shifts the cockpit air
supply control to full warm. The full warm air
supply automatically terminates 3 minutes after
activation. The lever may be cycled to restart thefull warm, 3minute period.
181

T.O. GR1F16CJ1
Under extremely humid conditions or after initialengine start, fog may form at the cockpit air outlets as
the cold air mixes with the moist cockpit atmosphere.
This condition can be eliminated by selecting MAN
and moving the TEMP knob toward WARM until thefog stops forming. In flight, while operating in AUTO,
the most rapid method of eliminating air outlet
fogging is by selecting the MAX position with the
DEFOG lever. Fog may form on the interior surface ofthe canopy as a result of moisture in the cockpit air
condensing on the cold surface. To warm the canopy
surface above the dewpoint and permit the cockpit air
to retain more moisture during cold weather
operation, the DEFOG lever should be placed in a
forward position and the TEMP knob positioned to
MAN WARM.
EQUIP HOT Caution Light

Defog Control Panel
C DF (Typical)

1

The EQUIP HOT caution light, located on the caution
light panel, illuminates when the avionic equipment
cooling air temperature/pressure is insufficient.
Degraded equipment performance and/or damage
can result from overheating. Therefore, when the
EQUIP HOT caution light illuminates, the electronicequipment should be turned off unless it is essential
for flight. Illumination of the EQUIP HOT caution
light automatically interrupts electrical power to theradar. Turning the radar to OFF in flight does not
close the radar cooling air shutoff valve.
A short duration or intermittent EQUIP HOT caution
light may occur when ground cooling air isdisconnected.
Cockpit Pressure Altimeter C

DF

The cockpit pressure altimeter, located on the right
auxiliary console outboard of the stick, is labeled
CABIN PRESS ALT.
CABIN PRESS Caution Light
The CABIN PRESS caution light, located on thecaution light panel, illuminates when the cockpit
pressure altitude is above 27,000 feet.
ANTIG SYSTEM
The antig system includes the ANTIG panel/valve,
the gsuit, and PBG equipment.
182

1. DEFOG Lever

GR1F-16CJ-1-0033X37

Figure 128.

The gsuit connector and TEST button are located on
the ANTIG panel at the aft end of the left console.
The ECS delivers cooled bleed air to the gsuit and tothe oxygen regulator as a control pressure for PBG.
Airflow is proportional to the positive g forces
sensed. If an ECS shutdown occurs, gsuit and PBG
protection are not available.
The system can be manually tested by depressing
the antig TEST button to inflate the gsuit and tocheck the PBG function. The system incorporates anautomatic pressure relief valve.

ELECTRICAL SYSTEM
Refer to figure 129. The electrical system consists ofa main ac power system, a standby ac power system,
an emergency ac power system, a dc power system,
an FLCS power supply, and provisions for externalac power.
MAIN AC POWER SYSTEM
AC power is normally supplied by a 60 kva main
generator located on and driven by the ADG. The
main generator supplies power to the overcurrent
sensing contactors and nonessential, essential, andemergency ac buses.

T.O. GR1F16CJ1
Overcurrent Sensing Contactors
The eight overcurrent sensing contactors protect
certain ac buses; stations 3, 5, and 7; and inlet
stations from overcurrent. The ELEC CAUTIONRESET button on the ELEC control panel is used toreset a tripped overcurrent sensing contactor on
nonessential ac bus No. 1 and the nacellenonessential ac bus. The overcurrent sensing
contactor may not remain reset if the fault persists.
The items with a nonresettable overcurrent sensing
contactor are the radar ac bus; stations 3, 5, and 7;
and left and right inlet stations.
STANDBY AC POWER SYSTEM
The standby ac power system consists of theessential and emergency ac buses and is powered (ifthe main generator is inoperative) by a 10 kva
standby generator which is located on and driven by
the ADG. The standby generator has power
available whenever the ADG is rotating and comeson line when the main generator is not supplying
power, as long as the MAIN PWR switch is in MAIN
PWR. The standby generator has an integral FLCS
PMG which supplies power to the four FLCS
branches. Refer to FLCS POWER SUPPLY, thissection, for further discussion of the FLCS PMG.
EMERGENCY AC POWER SYSTEM
If the main and standby generators fail, emergency ac
power is supplied automatically by a PX II 5 kva,
PX III 7 kva EPU generator driven by the EPU. The
system supplies power to the emergency ac buses.
The EPU generator has a PMG which supplies dc
power through an ac to dc converter to the four FLCS
branches. Refer to EMERGENCY POWER UNIT,
this section, for further discussion of the EPU.
DC POWER SYSTEM
DC power is supplied by ac to dc converters or by theaircraft battery. With the main generator operating,the ac to dc converters power emergency dc bus No.
1, battery bus No. 1, nonessential dc bus, nacelle dcbus, emergency dc bus No. 2, essential dc bus, andbattery bus No. 2. With the standby generator
operating, the ac to dc converters power emergency
dc bus No. 1, battery bus No. 1, emergency dc bus No.
2, essential dc bus, and battery bus No. 2. With the
EPU generator operating, the ac to dc converters
power emergency dc bus No. 1, battery bus No. 1,
emergency dc bus No. 2, and battery bus No. 2. With

the main, standby, or EPU generator operating, theaircraft battery is disconnected and charged by thebattery charger/control assembly. If all generators
fail, the aircraft battery is connected and powers
battery bus No. 1 and battery bus No. 2. The battery
buses are powered in all cases to provide a source ofpower to the FLCS power supply (if needed) and start
power to the EPU.
FLCS POWER SUPPLYThe primary FLCS power supply includes a
dedicated FLCS PMG and two dualchannel
converter regulators and four branch power supplies
within the FLCC.
The FLCS PMG is the primary power source for the
FLCS during normal operations. The FLCS PMG is
integral with the standby generator and generates
power whenever the ADG is rotating. The PMG hasfour outputs, one for each branch of the FLCS, andgenerates sufficient power to operate the FLCS at 40
percent rpm or greater.
Other FLCS power sources are the main generator,
the standby generator, the EPU generator, the EPU
PMG, and the aircraft battery.
Two converter/regulators, having two channels
each, provide a separate channel for each branch of
the FLCS. Both converter/regulators receive powerfrom the FLCS PMG, the aircraft battery, and if the
EPU is running, the EPU PMG. The branch A and B
converter/regulator also receives power from emer
gency dc bus No. 1, and the branch C and D
converter/regulator also receives power from emer
gency dc bus No. 2. Each converter/regulator
channel converts ac power from the FLCS PMG to dc,
selects the power source with the highest voltage(within limits), and provides dc power to therespective FLCC branch. Converter/regulator output voltages are regulated to prevent overvoltage to
the FLCS. The converter/regulators also provide
fault indications for display on the ELEC control
panel and provide test indications to the TEST
switch panel.
The aircraft battery can provide temporary emer
gency power to the FLCS. The length of time that theaircraft battery is able to power the FLCS is afunction of the state of charge. The FLCS
incorporates four latching relays which function toprevent depletion of the aircraft battery during
ground maintenance. The relays prevent the FLCC
from being connected to the aircraft battery until a
JFS start is initiated.

183

T.O. GR1F16CJ1Electrical Power Distribution Diagram
MONITOR UNIT

EXTERNAL PWR
CABLE RECEPTACLE

CSD

ADG

MAIN GENERATOR
PMG
MAIN POWER
CONTACTORSOVERCURRENTSENSINGCONTACTORSSTANDBY GENERATOR

F
W
D
A
C
LEFT
INLET

RIGHT
INLET

P
W
R
P
N
L

FLCS PMGNONESS AC
BUS NO. 1
STBY
ON

EPU CONTROLLERRADAR
AC BUSEPU GENERATOR
PMG

EPU
ON

AC TO DC CONVERTEREMERGENCYAC BUS NO. 1

A
B

AC TO DC CONVERTER
A
F
T
E
Q
U
I
P
B
A
Y
D
C
P
W
R
P
N
L

EMERGENCYDC BUS NO. 1

CONVERTER/ C
REGULATOR D

FLCC

AIRCRAFT BATTERYEPU ON(EXCEPT
DURING GND
TEST) ORSTBY ONNONESS DC
BUS

CONVERTER/
REGULATORBATTERY CHARGER, CONTROLASSEMBLY, AND FLCS RELAYSBATT BUS
NO. 1

GR1F-16CJ-1-0034-1B37

Figure 129.(Sheet 1)184

T.O. GR1F16CJ1OVERCURRENTSENSINGCONTACTORS

MAIN POWER
CONTACTORSSTATIONS 3, 5 & 7

NONESS AC
BUS NO. 2
STBY
ON

ESS AC
BUS
EPU
ON

EMERGENCYAC BUS NO. 2

R
I
G
H
T
S
T
R
A
K
E

NACELLE NONESS AC BUS
STBY
ON

A
C
P
W
R

NACELLE ESSAC BUSP
N
L

NACELLENONESSDC BUSNO. 2

AC TO DC CONVERTEREMERGENCYDC BUS NO. 2

S
T
R
A
K
E

EPU ON(EXCEPT
DURING GNDTEST)BATT BUS
NO. 2

ESS DCBUS

NACELLENONESS DC
BUS NO. 1

R
I
G
H
T

R
I
G
H
T

A
N C
A /
C D
E C
L
L P
E W
R
E
Q P
U N
I L
P
B
A
Y

LEGEND:
ELECTRICALMECHANICALRESETTABLE*
NONRESETTABLE*DC

Can be powered by EPU

PWRPowered thru NONESS AC
Bus No. 2

P
NL

Powered thru Fwd OCSCPowered thru Nacelle OCSC*OVERCURRENT SENSING CONTACTORS
GR1F-16CJ-1-0034-2A37

Figure 129.(Sheet 2)

185

T.O. GR1F16CJ1
EXTERNAL POWER PROVISIONSThe external power provisions include a standard
external power cable receptacle and a monitor unit
which is part of the airframe. The monitor unit
allows external power to be connected to the aircraft
buses if the phasing, voltage, and frequency of theexternal power are correct. When connected, theexternal power provides the same power as the main
generator.
ELECTRICAL SYSTEM CONTROLS AND INDICATORS
Refer to figure 130.
ELECTRICAL POWER DISTRIBUTION
Refer to figures 131 and 132.
ELECTRICAL SYSTEM NORMAL OPERATION
Prior to engine start, the MAIN PWR switch isplaced to BATT to permit a check of the aircraft
battery. The ELEC SYS, MAIN GEN, STBY GEN,and FLCS RLY lights come on. The FLCS RLY lightilluminates because the four FLCS relays are open
and the FLCC is not connected to the aircraft battery.
The FLCS PMG light is not illuminated since it
requires FLCS power. The ACFT BATT TO FLCS
light does not illuminate since the FLCS relays are
open. With the FLCS PWR TEST switch held in
TEST, the FLCS relays close but do not latch. The
FLCS PMG and the ACFT BATT TO FLCS lights
illuminate and the FLCS RLY light goes off. The
FLCS PWR lights on the TEST switch panel
illuminate, indicating that the power output of the
FLCC is good.

186

With the FLCS PWR TEST switch in NORM and theMAIN PWR switch positioned from BATT to MAIN
PWR, the lights do not change. If external power isconnected, the MAIN GEN light goes off.
If the FLCS PWR TEST switch is placed to TEST
with the MAIN PWR switch in MAIN PWR, theFLCS relays are latched closed. The FLCS RLY lightremains off when the FLCS PWR TEST switch isreturned to NORM. The aircraft battery is now
powering the FLCC and depletes more rapidly. To
eliminate the increased battery load, cycle the MAINPWR switch to BATT and back to MAIN PWR to open
the FLCS relays (FLCS RLY light illuminates).
When the JFS switch is moved to either start
position, the FLCS PMG light (and ACFT BATT TO
FLCS light if engine is started on battery power)
illuminates and the FLCS RLY light goes off,
indicating that the FLCS relays have closed.
During engine start, the FLCS PMG light (andACFT BATT TO FLCS light if engine is started onbattery power) goes off at approximately PW 229 40,
GE 129 44 percent engine rpm. The STBY GEN lightgoes off at approximately PW 229 55, GE 129 60
percent engine rpm; the MAIN GEN light goes off
approximately 10 seconds later if both generators
are operating normally. External power, if used, isdisconnected from the aircraft buses when the main
generator comes on line.
Anytime after selecting MAIN PWR, including inflight, the FLCS PWR TEST switch may be held
momentarily in TEST to check FLCC power output.
During the EPU test, the FLCS PWR lights come on
to indicate that EPU PMG power is available to the
FLCS.
During engine shutdown, the ELEC SYS caution
light and FLCS PMG, MAIN GEN, and STBY GEN
lights come on as the engine spools down. The ACFT
BATT TO FLCS light also illuminates.

T.O. GR1F16CJ1Electrical System Controls and Indicators
C DF (Typical)
12

3

1. MAIN PWR Switch
2. FLCS PMG Indicator Light
3. MAIN GEN Indicator Light
4. STBY GEN Indicator Light
5. EPU GEN Indicator Light
6. EPU PMG Indicator Light
7. ACFT BATT Indicator Lights
8. ELEC SYS Caution Light
9. ELEC CAUTION RESET Button
10. FLCS PWR TEST Switch
11. FLCS PWR Indicator Lights

4
5
6

9
11
10

7
8
GR1F-16CJ-1-0035X37

1.

CONTROL/INDICATORPOSITION/INDICATION

MAIN PWR Switch

MAIN PWR

Connects external power or the main genera
tor to the electrical system and enables
standby generator. Determines function of
FLCS PWR TEST switch. If ac power is notavailable, connects aircraft battery to thebattery buses

BATT

Connects aircraft battery to the battery
buses, disconnects main generator or exter
nal power, resets main generator, disables
standby generator, and determines functionof FLCS PWR TEST switch

OFF

In flight-disconnects main generator fromelectrical system and disables standby gener
ator

NOTE:
During ground operation, if
the MAIN PWR switch ismoved
d ffrom MAIN PWR tto
OFF without a delay of 1 second in BATT, the EPU does
not activate and electrical
power for braking, NWS,
hook,, and radios is lost.

FUNCTION

On ground-disconnects main generator orexternal power from aircraft electrical sys
tem and disables standby generator. Disconnects the aircraft battery from the battery
buses. Canopy operation is available afterengine shutdown
Figure 130.(Sheet 1)

187

T.O. GR1F16CJ1Electrical System Controls and Indicators
C DF (Typical)
2.

CONTROL/INDICATORPOSITION/INDICATIONFLCS PMG Indicator Light

FLCS PMG (amber)

FUNCTION
In flight-None of the FLCS branches arereceiving power from the FLCS PMG
On ground-FLCS PMG power is not avail
able at one or more FLCS branches. Light isdelayed 60 seconds after initial NLG WOW

3.

MAIN GEN Indicator Light

MAIN GEN (amber)

Indicates external power or main generator
not connected to one or both nonessential ac
buses

4.

STBY GEN Indicator LightSTBY GEN (amber)

Indicates standby generator power is notavailable

5.

EPU GEN Indicator LightEPU GEN (amber)

Indicates the EPU has been commanded on
but the EPU generator is not providing power
to both emergency ac buses. The light does
not function with the EPU switch in OFF(WOW) and the engine running6.

EPU PMG Indicator LightEPU PMG (amber)

Indicates the EPU has been commanded on
but EPU PMG power is not available to allbranches of the FLCS

7.

ACFT BATT Indicator
LightsFAIL (amber)

In flight-indicates aircraft battery failure(20V or less)On ground-indicates aircraft battery or bat
tery charger failure. Light is delayed 60
seconds after MLG WOW

TO FLCS (amber)

In flight-indicates battery bus power isgoing to one or more FLCS branches and volt
age is 25V or lessOn ground-indicates battery bus power isgoing to one or more FLCS branches

FLCS RLY (amber)

Figure 130.(Sheet 2)188

Indicates that voltage on one or more of thefour FLCC branches connected to the aircraft
battery is inadequate (below 20V) or that one
or more FLCC branches are not connected tothe battery

T.O. GR1F16CJ1Electrical System Controls and Indicators
C DF (Typical)
CONTROL/INDICATORPOSITION/INDICATIONFUNCTION

8.

ELEC SYS Caution LightELEC SYS (amber)

Illuminates in conjunction with any of theabove lights

9.

ELEC CAUTION RESET
ButtonPushResets resettable overcurrent protection
units and ELEC SYS caution light and clears
MASTER CAUTION light for future indica
tions. Resets main and standby generators

10.

FLCS PWR TEST Switch

TEST

When MAIN PWR switch is in:
MAIN PWR

BATT

Tests FLCC power Tests FLCC power
outputoutput on aircraft
battery

11.

FLCS PWR Indicator Lights

NORM

Normalposition. NA
Tests EPU PMG
power
availability
during
EPU/GEN
test on ground

MAINT

For maintenance use on the ground. Inopera
tive in flight

A, B, C, and D (green)

Illuminate to indicate proper power output of
FLCC during FLCS power tests

Figure 130.(Sheet 3)

189

T.O. GR1F16CJ1AC Power Distribution Diagram PX II
MAIN GENERATORNONESSENTIAL AC BUS NO. 2
Data Transfer Unit
Flood Console Lights
Flood Instrument Lights
Formation Lights
Fuel Pumps 1, 2, 4 & 5
Halon Heater
Inlet Strut Heater
Sta 4 & 6 - EO &Radar-Guided Weapons
Taxi LightSTBY GENERATOR

STBY
ON

ESSENTIAL AC BUS
EPU GENERATOR

Battery Charger Power
Fuel Pump No. 3
MFD's

EPU
ON

EMERGENCY AC BUS NO. 2
ADI
Anticollision LightAR Light (flood)Engine Ice Detector
Fire/Overheat Detect & Test

Fuel Quantity IndicatorGun
HSI
HUD/CTVSHYD PRESS Indicators
Landing LightC DF LIQUID OXYGEN
Quantity Indicator
NOZ POS Indicator
Oil Pressure Indicator
Position Lights
TACAN
Upfront Controls

GR1F-16CJ-1-1154-1A37

Figure 131.(Sheet 1)190

T.O. GR1F16CJ1OVERCURRENT SENSING CONTACTORSOVERCURRENT SENSING CONTACTORS
Sta 3

Sta 5

Sta 7

Sta 3

Sta 5

Sta 7

NacelleFCRFCR BUSSta 3, 5, & 7 - ECM, EO &Radar-Guided Weapons

Fwd

Left

Right

STBY
ON

FCR
LEFT
INLET

NACELLE NONESSENTIAL AC BUS

RIGHT
INLET

NONESSENTIAL AC BUS NO. 1

AIM-9/AIM-120
ECM Prcs

ECM Aft Xmit
FCC
GPS
Seat Adjustment
Total Temperature
Probe Heater
TWS

STBY
ON

STBY
ON

NACELLE ESSENTIAL AC BUS
Air Data Probe Heater(fuselage)
AIM-9/AIM-120
PFLD
EPU
ON

EMERGENCY AC BUS NO. 1
Air Data Probe Heater(nose)
Altimeter (ELECT)AOA IndicatorAOA Probe Heaters

CADC
DED
FUEL FLOW Indicator
INS
LEF's

LEGEND:

Low Energy Ignition
SystemPrimary Console Lights
Primary Instrument Lights
VVI

Can be powered by EPU

ELECTRICALMECHANICALRESETTABLENONRESETTABLE

Powered thru NONESS AC Bus No. 2
Powered thru Nacelle OCSCPowered thru Fwd OCSC
GR1F-16CJ-1-1154-2A37

Figure 131.(Sheet 2)

191

T.O. GR1F16CJ1DC Power Distribution Diagram PX IIEMERGENCY AC BUS NO. 1

AC TO DC CONVERTEREMERGENCY DC BUS NO. 1
C ALT REL Button
AOA Indexer
CADC Caution LightCIU
Cockpit Temp Control
DBU ON Warning Light(branches A & B)
EMER JETT Button

FLCS Reset Switches(branches A & B)
FLCS Power Source(branches A & B)
FLCS Warning Light(branches A & B)
C DF FLCS FAULT Caution
Light (branches A & B)

FUEL LOW Caution Lights
IFF
INS
INSTR MODE Select
LG DOWN Permission Button
LG WHEELS Down Lights
MAL & IND LTS Test/BRT/DIM Feature
MASTER ARM Switch

EPU ON(EXCEPT
DURING GND
TEST) ORSTBY ON

BATTERYNONESSENTIAL DC BUS
ECM Control
ECM Fwd Xmit
ECM Prcs
FCR
Fuel Pumps 1, 2 & 4
VHF Radio
NACELLE NONESS DC BUS NO. 1

MSL Step Switch
NWS/AR System
Probe Heat Monitor
PROBE HEAT Switch
Speedbrakes
Stick Trim
TACAN
Turn Needle
DR WPN REL Button

BATTERY BUS NO. 1
Antiskid/Chan 1
BrakesCANOPY Warning
Light
Cockpit Spotlights
Drag Chute
ENGINE Warning LightEPU
FLCS (branches A & B)

FTIT Indicator
HYD/OIL PRESS
Warning LightIntercom
JFS
LG Downlock Relays(WHEELS down lights)
LG Uplock/Downlock
Master Caution

MLG WOW(branches A & B)
NLG WOW(branches A & B)
OCSC Trip
RPM Indicator
UHF Radio
Utility LightVMS

ECM TBRFTWS
NACELLE NONESSENTIAL DC BUS NO. 2
AIBU
ECM Aft Xmit
EMSC

FUEL/OIL HOT Caution
Light (oil hot signal)
Nacelle Ejector Shutoff
GR1F-16CJ-1-1155-1A37

Figure 131.(Sheet 3)192

T.O. GR1F16CJ1EMERGENCY AC BUS NO. 2

AC TO DC CONVERTEREMERGENCY DC BUS NO. 2
ANTI ICE Switch
AR System Lights
Auto Fwd Fuel Trans
Autopilot
AVTR
CABIN PRESS CautionLight
CAMERA/GUN Trigger*Canopy Actuator
Chaff/Flare Dispensers
CIU
Cockpit Pressure Dump

DBU ON Warning Light(branches C & D)
EMER JETT Button
Engine Bleed Air Valves(close capability)
ENGINE FAULT Caution Light
EQUIP HOT Caution LightDR FLCS FAULT Caution
Light (branches C & D)
FLCS Power Source(branches C & D)

FLCS Reset Switch(branches C & D)
FLCS Warning Light(branches C & D)
FUEL/OIL HOT Caution
Light
Gun
HUD
ICP/IKP
ILS
INLET ICING CautionLight

EPU ON(EXCEPT
DURINGGND TEST)ESSENTIAL DC BUS

BATTERY BUS NO. 2
Altimeter (PNEU)
Antiskid/Chan 2 Brakes
DEC
Electrical Throttle
Position
ELEC SYS Caution LightEPU
EPU FUEL Quantity
Indicator

LANDING/TAXI/EXTLight Switches
LG Door Close
LG Hyd Isolation
LG UP-DN Command
MASTER ARM Switch
MFD Video Control
NUCLEAR CONSENT Switch
OXY LOW Caution Light
SEAT NOT ARMED
Caution LightSTORES CONFIG Caution
Light
Upfront Controls
C DF WPN REL Button

ENGINE FIRE Warning
Light
FLCS (branches C & D)
FUEL MASTER Switch
GCU's (main/standby)
Hook
LG Warning Light (handle)
LEF Asymmetry Brake
MLG WOW (branches
C & D)

NLG WOW (branches
C & D)OCSC Trip
OVERHEAT Caution
Light
Parking BrakeSAI
SEC Caution LightZEROIZE Switch

LEGEND:

Data Link
Fuel Pumps 3 & 5
Radar Altimeter
Stations 1, 2, 3, 3A, 4,
5, 6, 7, 7A, 8 & 9
ARM/Release
Tank Inerting

Can be powered by EPU

ELECTRICALMECHANICALPowered thru NONESS AC Bus No. 2
Powered thru Fwd OCSC

* Power available from the battery bus box with MAIN PWR switch OFF.

GR1F-16CJ-1-1155-2A37

Figure 131.(Sheet 4)

193

T.O. GR1F16CJ1AC Power Distribution Diagram PX III
MAIN GENERATORNONESSENTIAL AC BUS NO. 2
CFT Fuel Transfer Pumps
Data Transfer Unit
ECM Aft Transmitter
Flood Console Lights
Flood Instrument Lights
Formation Lights
Fuel Pumps 2 & 4
Halon Heater
Inlet Strut Heater
Sta 4 & 6 - EO &Radar-Guided Weapons
Taxi LightTISLSTBY GENERATOR

STBY
ON

ESSENTIAL AC BUS
EPU GENERATOR

Battery Charger Power
Fuel Pump No. 3
MFD's

EPU
ON

EMERGENCY AC BUS NO. 2
ADI
Anticollision LightAR Light (flood)
DTC
EGI
Engine Ice Detector

Fire/Overheat Detect & Test
Fuel Quantity IndicatorGun
HSI
HUD/CTVSHYD PRESS Indicators

Landing Light
NAV/Position Lights
NOZ POS Indicator
OBOGS Concentrator
Oil Pressure Indicator

GR1F-16CJ-1-2154-1X37

Figure 132.(Sheet 1)194

T.O. GR1F16CJ1OVERCURRENT SENSING CONTACTORSOVERCURRENT SENSING CONTACTORS
Sta 3

Sta 5

Sta 7

Sta 3

Sta 5

Sta 7

NacelleFCRFCR BUSSta 3, 5, & 7 - ECM, EO &Radar-Guided Weapons

Fwd

Left

Right

STBY
ON

FCR
LEFT
INLET

RIGHT
INLET

NACELLE NONESSENTIAL AC BUS
NONESSENTIAL AC BUS NO. 1

AIM-9/AIM-120
ECM Processor

ECM Fwd Transmitter
HMCS
Seat Adjustment
Total Temperature
Probe Heater
TWS

STBY
ON

STBY
ON

NACELLE ESSENTIAL AC BUS
Air Data Probe Heater(fuselage)
AIM-9/AIM-120
PFLD
EPU
ON

EMERGENCY AC BUS NO. 1
EGI (GPS)
FUEL FLOW Indicator
LEF's
MMC (Side A)

Altimeter (ELECT)AOA IndicatorAOA Probe Heaters
CADC

LEGEND:

Pitot Static Probe Heater
Primary Console Lights
Primary Instrument Lights
Upfront Controls
VVI

Can be powered by EPU

ELECTRICALMECHANICALRESETTABLENONRESETTABLE

Powered thru NONESS AC Bus No. 2
Powered thru Nacelle OCSCPowered thru Fwd OCSC
GR1F-16CJ-1-2154-2X37

Figure 132.(Sheet 2)

195

T.O. GR1F16CJ1DC Power Distribution Diagram PX III
EMERGENCY AC BUS NO. 1

AC TO DC CONVERTEREMERGENCY DC BUS NO. 1
C ALT REL Button
AOA Indexer
AVTR Switch
CADC Caution Light
Cockpit Temp Control
DBU ON Warning Light(branches A & B)
EGI
EMER JETT Button

FLCS Reset Switches(branches A & B)
FLCS Power Source(branches A & B)
FLCS Warning Light(branches A & B)
C DF FLCS FAULT Caution
Light (branches A & B)

FUEL LOW Caution Lights
IFF
LG DOWN Permission Button
LG WHEELS Down Lights
MAL & IND LTS Test/BRT/DIM Feature
MASTER ARM Switch
MMC (Side A)
MSL Step Switch

EPU ON(EXCEPT
DURING GND
TEST) ORSTBY ON

BATTERYNONESSENTIAL DC BUS
D ASIUECM Processor
AVTR Heater FCR
ECM Control Fuel Pumps 1, 2 & 4
ECM Fwd Transmitter

NWS/AR System
OBOGS Monitor
Probe Heat Monitor
PROBE HEAT Switch
Speedbrakes
Stick Trim
TACAN
VHF Radio
DR WPN REL Button

BATTERY BUS NO. 1
Antiskid/Chan 1
BrakesCANOPY Warning
Light
Drag Chute
ENGINE Warning LightEPU

NACELLE NONESS DC BUS NO. 1

FTIT Indicator
HYD/OIL PRESS
Warning LightIntercom
JFS
LG Downlock Relays(WHEELS down lights)
LG Uplock/Downlock
Master Caution

MLG WOW(branches A & B)
NLG WOW(branches A & B)
OCSC Trip
RPM Indicator
UHF Radio
Utility LightVMS

MFD Heater
TWS
NACELLE NONESSENTIAL DC BUS NO. 2
ECM Aft Transmitter
MAX POWER Switch
Nacelle Ejector Shutoff
TGP (right inlet power and control)
TISL

Figure 132.(Sheet 3)196

GR1F-16CJ-1-2155-1X37

T.O. GR1F16CJ1EMERGENCY AC BUS NO. 2

AC TO DC CONVERTEREMERGENCY DC BUS NO. 2
ANTI ICE Switch
Antiskid/Chan 2 Brakes
AR System Lights
Auto Fwd Fuel Trans
Autopilot
AVTR
CABIN PRESS CautionLight
CAMERA/GUN Trigger*Canopy Actuator
Chaff/Flare Dispensers
Cockpit Pressure Dump

Data Transfer UnitDBU ON Warning Light(branches C & D)
EMER JETT Button
Engine Bleed Air Valves(close capability)
ENGINE FAULT Caution Light
EQUIP HOT Caution LightDR FLCS FAULT Caution
Light (branches C & D)
FLCS Power Source(branches C & D)

FLCS Reset Switch(branches C & D)
FLCS Warning Light(branches C & D)
Gun
HUD
ICP/IKP
ILS
INLET ICING CautionLight

EPU ON(EXCEPT
DURINGGND TEST)ESSENTIAL DC BUS

BATTERY BUS NO. 2
Altimeter (PNEU)
Antiskid/Chan 2 Brakes
CSFDR
EDU
ELEC SYS Caution LightEPU
EPU FUEL Quantity
Indicator

LANDING/TAXI/EXTLight Switches
LG Door Close
LG Hyd Isolation
LG UP-DN Command
MASTER ARM Switch
MFD Video Control
MMC (side B)
NUCLEAR CONSENT Switch
OBOGS Caution LightOXY LOW Warning Light
SEAT NOT ARMED
Caution LightSTORES CONFIG Caution
Light
Upfront Controls
C DF WPN REL Button

ENGINE FIRE Warning
Light
FUEL MASTER Switch
GCU's (main/standby)
Hook
LEF Asymmetry Brake
LG Warning Light (handle)
MLG WOW (branches
C & D)

NLG WOW (branches
C & D)OCSC Trip
OVERHEAT Caution
Light
Parking BrakeSAIZEROIZE Switch

LEGEND:

D Aft Seat Hud Monitor
Data Link
Fuel Pumps 3 & 5
NVP (left inlet
power and control)
Radar Altimeter
Stations 1, 2, 3, 3A, 4,
5, 6, 7, 7A, 8 & 9
ARM/Release
Tank Inerting

Can be powered by EPU

ELECTRICALMECHANICALPowered thru NONESS AC Bus No. 2
Powered thru Fwd OCSC

* Power available from the battery bus box with MAIN PWR switch OFF.

GR1F-16CJ-1-2155-2X37

Figure 132.(Sheet 4)

Change 1197

T.O. GR1F16CJ1HYDRAULIC SYSTEMHYD PRESS INDICATORS AND WARNING LIGHT
Refer to figure 134.

Refer to figure 133. Hydraulic pressure is suppliedby 3000 psi hydraulic systems designated as systems
A and B. The systems are powered by two
independent enginedriven pumps located on the
ADG. Each system has a reservoir to store hydraulicfluid. The reservoirs are pressurized by theirrespective hydraulic system to insure positive
pressure at the pump. For hydraulic system cooling,refer to ENGINE FUEL SUPPLY SYSTEM, thissection.
Both systems operate simultaneously to supply
hydraulic power for the primary flight controls andLEF's. If one of the systems should fail, theremaining system provides sufficient hydraulicpressure; however, the maximum actuation rate of
the FLCS is reduced. System A also supplies power
to the FFP and the speedbrakes. All remaining
utility functions, consisting of the gun and gun purgedoor, AR system, LG, brakes, and NWS, are suppliedby system B. System B also charges the brake/JFS
accumulators (which provide start power for the JFS
and backup pressure for the brakes), provided theengine is rotating at a minimum of 12 percent rpm.
System B takes between 40 seconds (hot ambientconditions) and 60 seconds (cold ambient conditions)
to recharge the brake/JFS accumulators.
The LG can be extended pneumatically in the eventof hydraulic system B failure. Should both hydraulicsystems fail, a third hydraulic pump located on the
EPU automatically provides hydraulic pressure tosystem A. Refer to EMERGENCY POWER UNIT
(EPU), this section, for a further discussion of the
EPU.
Each hydraulic system has an FLCS accumulator
which is isolated from the main system by checkvalves. These FLCS accumulators serve a dual
function. If demand exceeds the pump maximum
flow rate during rapid control surface movement, theFLCS accumulators provide additional hydraulicpressure. Also, if both hydraulic systems fail, theFLCS accumulators provide adequate hydraulicpressure to the flight controls while the EPU comesup to speed. Refer to SERVICING DIAGRAM, thissection, for servicing/specifications information.

198HYD PRESS Indicators
The HYD PRESS indicators, one for system A and
one for system B, are located on the right auxiliary
console. The indicators are powered by emergency ac
bus No. 2.
HYD/OIL PRESS Warning LightA HYD/OIL PRESS warning light, located on theright glareshield, comes on when hydraulic system
A or B pressure drops below 1000 psi or when engine
oil pressure drops below 10 ("2) psi. For the oilpressure function only, there is a 30second timedelay in the light circuit to minimize transient lights
during negative g maneuvers. The light is poweredby battery bus No. 1.

EMERGENCY POWER UNIT (EPU)
Refer to figure 135. The EPU is a selfcontained
system which simultaneously provides emergency
hydraulic pressure to system A and emergency
electrical power. The EPU automatically activateswhen both main and standby generators fail or whenboth hydraulic system pressures fall below 1000 psi.
The EPU may be operated manually regardless offailure conditions.
The EPU requires dc power from either battery bus
No. 1 or No. 2 for automatic or manual activation.
When the EPU is operating, the emergency ac and dc
buses are powered by the EPU generator. To reduce
electrical loads, the nonessential and essential dc
buses are unpowered whenever the EPU is operating(except when activated for ground test using the
EPU/GEN test switch).
When operating, the EPU augments hydraulicsystem A as required. If the normal system A
hydraulic pump fails, the EPU is the only source ofsystem A pressure.
The EPU uses engine bleed air and/or hydrazine tooperate. Normally, engine bleed air is used tomaintain operating speed. When bleed air isinsufficient, hydrazine augmentation automaticallyoccurs. Hydrazine is always used when the EPU iscommanded to start except when activated during
ground test using the EPU/GEN test switch.

T.O. GR1F16CJ1
On system command, hydrazine is forced by nitrogenpressure into a decomposition chamber. The gaseous
products of the reaction spin the turbine/gearbox
which then powers the EPU generator and hydraulicpump. Hydrazine exhaust is vented overboard on thelower inboard side of the right strake and consists
primarily of nitrogen, hydrogen, ammonia, andwater. The temperature of exhaust gases can reach
1600_F and will ignite in the presence of a flame. The
exhaust gases have an ammonia odor, are irritating
to the nose and eyes, and should be avoided to the
maximum extent possible.
EPU CONTROLS AND INDICATORS
Refer to figure 136.
EPU Ground Safety Switch
The ground safety switch, located on the right side ofthe engine inlet, is used to disable the EPU on the
ground. With the EPU safety pin installed, the EPU
does not operate.
EPU Switch
The EPU switch, located on the EPU control panel, isa threeposition toggle switch.
Functions are:
D OFF
S Prevents or terminates EPU operation on the

ground (WOW).

S Does not prevent or terminate EPU operation inflight for main and standby generator failures if
switch was cycled or placed to NORM at anytime
since takeoff (since WOW).

S Prevents EPU operation in flight if switch has

remained in the OFF position since takeoff (since
WOW).

S Terminates EPU operation in flight except duringmain and standby generator failures.

D NORM- The system is armed for automatic
operation except during engine shutdown on the
ground. With WOW and throttle in OFF, the EPU
does not activate when the main and standby
generators drop off line.

D ON - Commands EPU to run regardless of failure

conditions. Operation will cease when switch is
positioned to OFF except for main and standby
generator failures in flight.

The switch has a split guard; the top half can be raised
to move the switch to ON, and the bottom half can beraised to move the switch to OFF. When both sectionsof the guard are down, the switch is retained in the
NORM position.
EPU Run Light
The EPU run light, located on the EPU control panel,
illuminates when the EPU turbine speed is within theproper range and the EPUdriven hydraulic pumpdischarge pressure is above 2000 psi.
HYDRAZN Light
The HYDRAZN light, located on the EPU control
panel, illuminates when the EPU is commanding
hydrazine for operation (whether hydrazine isavailable or not) or if a primary speed control failure
has occurred.
AIR Light
The AIR light, located on the EPU control panel,
illuminates whenever the EPU has been commanded
to run with the EPU safety pin removed. It remains
on even when the EPU is augmented by hydrazine.
EPU/GEN Test Switch
The EPU/GEN test switch, located on the TEST
switch panel, has positions of OFF and EPU/GEN.
The switch is springloaded to the OFF position. It
provides a means to test the EPU generator and EPU
PMG output to FLCS on the ground without using
hydrazine.
EPU FUEL Quantity Indicator C

DF

Refer to figure 13. The EPU FUEL quantity
indicator, located on the right auxiliary console, isgraduated 0100 and indicates the percent ofhydrazine remaining. The indicator is powered bybattery bus No. 2.
Hydrazine Leak Detector
The hydrazine leak detector is a silicone base,mustard yellow disc visible through access door 3208.
The viewing area is black on one half to provide
contrast with the mustard yellow disc. The mustard
yellow turns purple/black in the presence of
hydrazine and/or its vapors, indicating a leak in the
EPU and/or fuel tank system.

199

T.O. GR1F16CJ1

Hydraulic Power Supply(Typical)
SYSTEM A

B
P

THERMAL
BYPASS
FUEL/OILHEAT EXCHANGERFILTERFLCS
ACCUMULATOR

LEF'SENGINEDRIVENHYDRAULICPUMPRUDDER

HORIZONTAL
TAILS

EPU
HYDRAULIC
PUMP

FLAPERONSSPEEDBRAKESFILTERFFPS
LG SYSTEMT

HYD/OILNWSPRESSWARNING LIGHT

BRAKE/JFSACCUMULATORSDRAG CHUTE
SYSTEM

GUN SYSTEMHYD PRESS
INDICATORAR SYSTEMHYDRAULICRESERVOIRDUMPAND
RELIEFMANIFOLDBRAKE SYSTEMJFS SYSTEMDUMP ANDRELIEF VALVERESERVOIRPRESSURIZATIONACCUMULATORRESERVOIRPRESSURIZATIONACCUMULATOR

GR1F-16CJ-1-0037-1X37

Figure 133.(Sheet 1)1100

T.O. GR1F16CJ1B
FILTER P

FLCS
ACCUMULATORSYSTEM B
ENGINEDRIVENHYDRAULIC
PUMPFILTERTS
LG DOORSUPLOCKACTUATEDLEGEND:
ELECTRICAL

DRAG CHUTE
ACCUMULATORISOLATION VALVEHYD PRESS
INDICATORFILTERBLEED VALVE

HYD/OIL

FILL CONNECTORPRESSQUICK DISCONNECTWARNING LIGHT

HYDRAULICRESERVOIRCHECK VALVESPRESSURE SWITCHT

PRESSURE TRANSMITTER

B
P

BYPASSPRESSURE
RETURN
SUPPLYDUMP ANDRELIEF VALVE
GR1F-16CJ-1-0037-2X37

Figure 133.(Sheet 2)

1101

T.O. GR1F16CJ1

HYD PRESS Indicators and Warning Light(Typical)
1

C DF

1

DR

2

3

1. HYD/OIL PRESS Warning Light (Red)
2. System A HYD PRESS Indicator
3. System B HYD PRESS Indicator2

3
GR1F16CJ10036X37

Figure 134.

EPU Fired Indicator
The EPU fired indicator is located next to the EPU
ground safety switch on the right side of the engine
inlet. Normally, the indicator displays a gray andblack disc. If the EPU has been activated, theindicator displays six equally spaced black and white
triangles.

EPU OPERATIONThe EPU is designed to operate automatically formain and standby generator failure, dual hydraulicsystem failure, PTO shaft or ADG failure, and engine
flameout or if the engine is shut down in flight. The
EPU can also be activated manually. After receiving
any start command, the EPU requires approximately
2 seconds to come up to speed.

1102

EPU startup may not be audible. Once operating,however, the EPU may be heard but does not soundthe same as during the EPU ground check. A lack ofsound during EPU startup does not indicate lack of
EPU operation which must be confirmed bymonitoring the EPU run light. EPU rpm is controlled
by three speed controls. The primary and secondary
speed controls are based on EPU rpm. The tertiary
speed control is based on EPU PMG frequency.
When the EPU is operating, engine thrust settings
should be maintained to prevent using hydrazine.
This normally requires a minimum of PW 229 7580,
GE 129 8290 percent rpm depending on pressure
altitude.
If the engine fails, hydrazine alone is used to power
the EPU. With hydrazine only, operating time of thesystem is approximately 10 minutes under normalload requirements. Increased flight control move
ment reduces this operating time. When the EPU isthe sole source of hydraulic power, EPU loss resultsin loss of aircraft control.

T.O. GR1F16CJ1Emergency Power Unit Schematic (Typical)
NOTE:
Shown in augment mode with mainand standby generators failed.

FROM HYD SYS A RSVRTO HYDSYS A
CLOSED FOR
PRESSURE > 2000 PSIEPU RUN LIGHT(BLEED AIR
ONLY)
EPU

SPEED
MONITORTO CONVERTER/
REGULATORS

ELEC
SYS

TEST RUNCOMMANDHYDRAULIC
PUMP

P
M
G

TO EMERGENCYAC BUSES
MAIN GENERATOR
FAIL INDICATION

5 KVA
PX III 7 KVA
GENERATOR(AUGMENT)PX II

AC POWEREPU FUELQUANTITYCONTROL VALVESOVERSPEED SHUTOFF COMMAND(TERTIARY CONTROL)
GENERATORCONTROL UNIT

STANDBY GENERATOR
FAIL INDICATIONBLEED AIRSOLENOIDVALVE

EPU GENERATOR FAIL INDICATIONEPU PMG FAIL INDICATION FROM CONVERTER/REGULATORS
VALVE
ASSY

PRESSURE
INDICATOREPU/GEN TESTSWITCHHIGH PRESSUREBLEED AIRHYDRAZINE FUELNITROGENEPU COMMAND INDICATIONHYD SYS A

RUN LIGHT OFF
FOR OVER OR
UNDERSPEED
CONDITIONRUNCOMMAND

OPEN WITHWOW &THROTTLE
BELOW IDLE
AUTO ARM/
DISABLECLOSED ON
START, MAIN
GEN ON LINEEPU GROUNDSAFETY SWITCHEPU
CONTROLLER

HYD SYS B

CLOSED FOR
PRESSURE < 1000 PSI
CLOSED IF
STANDBY GEN
OFF LINE

CLOSED IF
MAIN GEN
OFF LINE

MAIN
GENERATORSTANDBY
GENERATOREPU TEST RUN COMMAND

AUGMENTMODE

BLEED AIR
MODE ONLYEPU START UP AND CONTROL POWER
LEGEND:BLEED AIRHYDRAZINE FUEL

28 VDC PWR

HYDRAULIC FLUID SUPPLYOPEN IF EPU
SAFETY PININSTALLEDHYDRAULIC PRESSUREGASEOUS NITROGEN

ELECTRICALMECHANICALFILL
CONNECTORQUICKDISCONNECT
GR1F-16CJ-1-0038A37

Figure 135.

1103

T.O. GR1F16CJ1EPU Control Panel
C DF (Typical)

LANDING GEAR CONTROLS AND INDICATORS
Refer to figure 137. The LG control panel is locatedon the left auxiliary console.

Landing Gear Handle
1
2
3

4

1. HYDRAZN Light (Amber)
2. AIR Light (Amber)
3. EPU Run Light (Green)
4. EPU Switch
GR1F-16CJ-1-0039X37

Figure 136.

LANDING GEAR (LG) SYSTEM

The LG handle, located on the LG control panel has
a wheelshaped grip. Movement of the handle
operates electrical switches (powered by emergency
dc bus No. 2) to command LG retraction or extension.
A warning light in the LG handle, powered by battery
bus No. 2 illuminates when the LG and doors are intransit or have failed to lock in the commanded
position. The warning light also illuminates when all
LG's are not down and locked, airspeed is less than
190 knots, altitude is less than 10,000 feet, and rateof descent is greater than 250 feet per minute. The
handle is locked in the DN position when the aircraft
is on the ground (weight on wheels). In flight, a signal
from the left MLG WOW switch automatically
activates a solenoid which unlocks the handle,
allowing movement to the UP position. The handle islocked in the UP position to prevent LG extension
during high g maneuvers.

Landing Gear Handle Down Permission Button

The LG system is normally operated by hydraulicsystem B. The NLG is extended and retracted byhydraulic pressure. The MLG's are retracted
hydraulically but are extended by freefall assisted byairloads. All the LG doors are hydraulically activatedwith electrical sequencing during retraction andmechanical sequencing during extension. If hydrau
lic system B fails, the LG may be extended
pneumatically.

The LG handle down permission button, located on the
LG handle, unlocks the handle electrically to permit
movement to the DN position. The button energizes anelectrical solenoid which releases the springactuatedhandle lock. The button must be depressed before
downward force is applied to the LG handle. The
electrical solenoid may not unlock the handle while
any appreciable downward force is applied.

MAIN LANDING GEAR (MLG)

DN LOCK REL Button

The two MLG's are independent of each other andretract forward with a mechanical wheel twist into two
separate wheel wells. Each MLG wheel is equippedwith three fusible (thermal pressure relief) plugs.
NOSE LANDING GEAR (NLG)
The NLG retracts aft with a 90degree mechanicalwheel twist into the wheel well. A torque arm
quickdisconnect is provided so that the nosewheel
can be turned beyond the steerable range for towing.
1104C DF The DN LOCK REL button, located on the LG
control panel, when depressed, mechanically unlocks
the springactuated handle lock if the electricalsolenoid should fail or not be powered. It overrides allelectrical LG control signals. Depressing this button
and raising the LG handle on the ground retracts the
LG. The downlock release button may not unlock the
LG handle while any appreciable downward force isapplied. DR For DN LOCK REL button differences,refer to F16D AIRCRAFT, this section.

T.O. GR1F16CJ1Landing Gear Control Panel (Typical)
3

2
1
C

DF

4
5

16

6

15

7

14
13

8

12

11

10

9

2

3

1

DR

5
16
6
7

17
1. EMER STORES JETTISON Button(Covered)
2. WHEELS Down Lights (Green)
3. HOOK Switch (Lever Lock)
4. ANTI-SKID Switch
5. DN LOCK REL Button
6. LG Handle Down Permission Button
7. LG Handle
8. LG Handle Warning Light (Red)
9. LANDING TAXI LIGHTS Switch
10. HORN SILENCER Button
11. ALT GEAR Handle
12. ALT GEAR Reset Button
13. SPEED BRAKE Position Indicator
14. STORES CONFIG Switch
15. BRAKES Channel Switch
16. GND JETT ENABLE Switch (Lever Lock)
17. ALT FLAPS Switch

10
8

13

12
11

GR1F-16CJ-1-0040A37

Figure 137.

1105

T.O. GR1F16CJ1
ALT GEAR Handle

WHEELS Down Lights

The ALT GEAR handle, located just outboard of andbelow the LG control panel, is used to extend the LG
if normal extension is not possible. Pulling the ALT
GEAR handle supplies pneumatic pressure to open all
LG doors, extend the NLG, and shut off the LG selector
hydraulic valve. The LG/hook emergency pneumatic
bottle is also used to lower the hook and contains
sufficient pneumatic pressure for one LG extension
and to hold the hook down. The bottle cannot berecharged in flight. Since pneumatic pressure is
reduced by expansion as the actuators extend, lessthan the normal extending force is available.

The three green WHEELS down lights, located on
the LG control panel, are arranged on the silhouetteof the aircraft. When any LG is down, its respective
light is on. A safe up and locked LG condition isindicated when all three of the lights and the LG
handle warning light are off. The lights are poweredby battery bus No. 1 (to energize downlock relays)
and emergency dc bus No. 1 (to illuminate lights).

An LG reset button, located on the ALT GEAR handle,
provides a means of retracting the LG after an alternate
extension if system B hydraulic pressure is available.
LG Warning Horn
The LG warning horn is an intermittent fixed volume
signal which sounds in the headset when the NLG or
MLG is not down and locked and all the following
conditions exist:

Landing
Switches

Gear

WeightonWheels

The LG WOW switches, located on both MLG's and on
the NLG, operate as a function of LG strut extension
to allow or terminate various system functions.
Refer to figure 138 for a list of systems affected by the
WOW switches and symptoms of WOW switch failure.
LANDING GEAR OPERATION
Movement of the LG handle to the UP position causes
the following events:

D Airspeed is below 190 knots.

D LG handle warning light illuminates.

D Pressure altitude is less than 10,000 feet.

D LG unlocks and retracts.

D Rate of descent is greater than 250 fpm.

D Three WHEELS down lights go off.

HORN SILENCER Button

D MLG wheel spin is stopped.

The HORN SILENCER button is located on the LG
control panel. Depressing the button silences the LG
warning horn. If the warning condition is eliminated,the horn resets. If it is not eliminated, subsequent LG
audio warnings do not occur.

(WOW)

D LG doors close and lock.
D LG handle warning light goes off.
D Hydraulic pressure removed from LG.
D FLCS switches to cruise gains.

TO/LDG CONFIG Warning LightD TEF's retract to streamlined position.

The TO/LDG CONFIG warning light, located on theright glareshield, illuminates in flight wheneverpressure altitude is less than 10,000 feet, airspeed isless than 190 knots, rate of descent is greater than
250 fpm, and either of the following conditions exists:

D Electrical power is removed from brake channel 1.

D TEF's not full down.

D LG doors and LG unlock, extend, and lock into place.

D NLG

D Three WHEELS down lights illuminate.

or either MLG not down and locked(accompanied by LG warning horn).

The TO/LDG CONFIG warning light illuminates on
the ground if TEF's are not full down.
With TEF's full down, rapid reversals of rollcommand inputs may cause the TO/LDG CONFIG
warning light to momentarily illuminate if thealtitude, airspeed, and rate of descent conditions
outlined above are met or WOW.
1106

Movement of the LG handle to the DN position causes
the following events:
D LG handle warning light illuminates.

D LG handle warning light goes off.
D TEF's extend.
D FLCS switches to takeoff and landing gains.
D PW 229 LG DN nozzle scheduling is activated.
D Speedbrakes close to 43 degrees if not overridden.
D Electrical power is supplied to brake channel 1.

T.O. GR1F16CJ1LG WOW Switch
RIGHT MLGSYSTEMS

Aircraft Battery

FCR

Air Data Probe

FLCC

Altimeter (ELECT)LG Warning

AOA Probes

Pitot Probe

Brakes/Antiskid

Probe Heat Monitor

ECS
Engine Controls

SMS

EPU

C
VMS

Fails to Ground Position in Flight

Fails to InFlight Position on Ground

FCR does not transmit

FCR transmits

Stores cannot be emergency jettisoned with LG down unless
GND JETT ENABLE switch is in ENABLE

Stores can be emergency jettisoned with GND JETT ENABLE switch
in OFFC VMS is inoperative
C VMS is operative unless INHIBIT is selected

Brakes can be applied before touchdown if toe brakes are
depressed

With simultaneous failure of left and right MLG WOW switches
and ANTISKID switch in ANTISKID, toe brakes are inoperative
when groundspeed is less than 20 knots

ANTISKID switch holds in PARKING BRAKE with throttle in OFF to
IDLE range

With simultaneous failure of left and right MLG WOW switches,
ANTISKID switch must be held in PARKING BRAKE to operateparking brake

With simultaneous failure of left and right MLG WOW switches,
ACFT BATT FAIL light indicates aircraft battery failure (voltage 20Vor less) or battery charger failure

ACFT BATT FAIL light indicates aircraft battery failure only
EPU is commanded on during engine shutdown; operation cannot
be terminated with the EPU switchCLG
and low speed warning tones are inoperative
Probe heat monitor is inoperative unless TEST or PROBE HEAT isselected

Probe heat monitor is operative

All probe heaters except total temperature are on
FLCS BIT cannot be initiatedFigure 138.(Sheet 1)

1107

T.O. GR1F16CJ1LG WOW Switch
LEFT MLGSYSTEMS

Aircraft Battery

Ground Test Panel (fuel pump lights)

AOA Probe (right)

JFS Ground Cutout

Brakes/Antiskid

LG Handle

Chaff/Flare Dispenser

LG Warning

EPU

Total Temperature Probe

FLCC

D
VMS

Fails to Ground Position in Flight

Fails to InFlight Position on Ground

Chaff/flare dispenser is inoperative

Chaff/flare dispenser is operative
AOA probe heater (right) is on

JFS shuts down automatically during engine start

JFS does not shut down automatically during engine start

LG UP position cannot be selected unless DN LOCK REL button isdepressed

LG UP position can be selected without DN LOCK REL button
depressed

TO/LDG CONFIG warning light is on with TEF's not down

TO/LDG CONFIG warning light is off with TEF's up

Total temperature probe heater is inoperative

Total temperature probe heater is on

DVMS
is inoperativeDVMS
is operative unless INHIBIT is selected

Brakes can be applied before touchdown if toe brakes are
depressed

With simultaneous failure of right and left MLG WOW switches
and ANTISKID switch in ANTISKID, toe brakes are inoperative
when groundspeed is less than 20 knots

ANTISKID switch holds in PARKING BRAKE with throttle in OFF to
IDLE range

With simultaneous failure of right and left MLG WOW switches,
ANTISKID switch must be held in PARKING BRAKE to operateparking brake
Fuel pump lights on external ground test panel are inoperative

With simultaneous failure of right and left MLG WOW switches,
ACFT BATT FAIL light indicates aircraft battery failure (voltage 20Vor less) or battery charger failure

ACFT BATT FAIL light indicates aircraft battery failure onlyDLGand low speed warning tones are inoperative

EPU is commanded on during engine shutdown; operation cannot
be terminated with the EPU switch
FLCS BIT cannot be initiatedFigure 138.(Sheet 2)

1108

T.O. GR1F16CJ1LG WOW Switch
NLG

SYSTEMSAir Data Probe

FLCS PowerAOA Probe (Left)

NWS
C

AR
C

Pitot Probe

FLCP

Speedbrakes

Fails to Ground Position in Flight

Fails to InFlight Position on GroundAir data, pitot, and left AOA probe heaters are on
AOA indicator displays zero degrees
A/R
DISC button is inoperativeC

A/R
DISC button is operativeC
FLCS fails BIT

NWS can be engaged and follows rudder inputs with NLG down

NWS is inoperative

Speedbrakes are not limited to 43 degrees with right MLG down
and locked

Speedbrakes do not remain open more than 43 degrees

ACFT BATT TO FLCS light indicates aircraft battery is supplying
power to one or more FLCS branches

ACFT BATT TO FLCS light indicates aircraft battery is supplying
power to one or more FLCS branches (bus voltage 25 vdc or less)

FLCS PMG light indicates FLCS PMG power is not available at one
or more FLCS branches

FLCS PMG light indicates the FLCS PMG is not supplying power to
any FLCS branches

Figure 138.(Sheet 3)

1109

T.O. GR1F16CJ1NOSEWHEEL STEERING (NWS) SYSTEMThe NWS is electrically controlled using dc bus No. 1
power and is hydraulically operated using system B
pressure. Steering signals are provided through therudder pedals. Should NWS be engaged with therudder pedals displaced, the nosewheel drives to therudder pedal commanded position. NWS is limited to
32 degrees in each direction; however, turn radius canbe reduced by using inside brake. NWS is
automatically disengaged when the NLG strut is
extended. NWS is not available following analternate LG extension and may not be available
anytime the nosewheel WHEELS down light is notilluminated.
NWS CONTROLS AND INDICATORS
Refer to figure 13.
NWS A/R DISC MSL STEP Button
The NWS A/R DISC MSL STEP button, located on theoutboard side of the stick, is used to engage or
disengage NWS when the aircraft is on the ground.
Once depressed, NWS is engaged and the button may
be released. If the button is held depressed,
continuous NWS is provided.
NWS Light
The NWS light, the center element of the AR/NWS
status indicator located on the top of the glareshield,
illuminates green when NWS is engaged. NWS does
not operate even though the NWS light is illuminatedwhen the NWS FAIL caution light is on or whensystem B hydraulic pressure is unavailable. On the
ground, NWS continues to operate with the AIR
REFUEL switch in OPEN even though the NWS lightis off.
NWS FAIL Caution Light
The NWS FAIL caution light, located on the caution
light panel, illuminates when a failure in the NWS
system has caused electrical power to be switched off.
NOSEWHEEL STEERING (NWS) SYSTEM

D

Either cockpit can control the NWS. Control isaccomplished by means of an NWS control button/in
dicator located in both cockpits just aft of the stick.
Depressing the button/indicator transfers control of
the NWS; the indicator illuminates green in the
cockpit which has control. Control then remains inthat cockpit even if the engine is shut down and/or
electrical power is removed. NWS can only be
1110Change 1

engaged and operated from the cockpit which has
control. When NWS is selected, the NWS lightilluminates green in both cockpits. The aft stick
paddle switch allows immediate override of theforward cockpit if the STICK CONTROL switch on
the TEST switch panel is in AFT; the NWS control
button/indicator in the aft cockpit and OVRD lightilluminates. The front cockpit paddle switch cannot
override the aft cockpit regardless of the position of
the STICK CONTROL switch. With the AIR
REFUEL switch in OPEN, NWS can be engaged ordisengaged from either cockpit without using thepaddle switch regardless of which cockpit has control.

WHEEL BRAKE SYSTEM
Each MLG wheel is equipped with a hydraulicallypowered multiple disc brake. The brakes areelectrically controlled by conventional toe brake
pedals. The amount of braking gradually increases as
pedal pressure is applied. A parking brake is also
provided. An antiskid system protects against blown
tires and is only available when using toe brakes.
D Brakes may be applied singly or simultaneously
from the forward or rear cockpits. The brake signals
from both cockpits are additive so that the total signal
to the brakes is the sum of the pedal forces from both
cockpits.

Brake hydraulic power is supplied by system B. If
system B fails or the engine is operating at less than
12 percent rpm, the toe brakes and parking brake areavailable until the brake/JFS accumulators deplete.
Continuous use of the toe brakes, even with theparking brake set, depletes accumulator fluid andcauses loss of all braking capability after approxi
mately 75 seconds (brake/JFS accumulators initiallyfully charged). When holding the aircraft stationary,use of the parking brake is preferred since brake/JFS
accumulator fluid is not depleted.
TOE BRAKE SYSTEM
Refer to figures 139 and 140. The toe brakes use
electrical power from the FLCC and CHAN 1 and
CHAN 2 dc power sources. The brake pedals require
FLCC power to operate. The pedal signals aresupplied to the b2t brake control/antiskid assembly
LESS b2t brake control box which, in turn, uses both
CHAN 1 and CHAN 2 dc power sources to operate
valves for controlling hydraulic pressure to thebrakes. CHAN 1 and CHAN 2 are powered by battery
buses No. 1 and No. 2, respectively. PX III In addition,
there is an alternate power source for CHAN 2 on
emergency dc bus No. 2.

T.O. GR1F16CJ1

Toe Brake System
BRAKE
CHAN 1FLCC
BRANCHESTOE BRAKES

C

BRAKE
CHAN 2

A

L

D

B

R

Regardless of which channel is selected, hydraulicpressure to three of the six pistons in each brake iscontrolled by electrical power from one dc powersource and pressure to the other three pistons of eachbrake is controlled by electrical power from anotherdc power source. A loss of one dc power source when
CHAN 1 is selected results in degraded brake
operation (only onehalf of the pistons are powered
and significantly more brake pedal force than normal
is required to stop). Due to redundancy features,
selecting CHAN 2 may restore full braking. If all dc
power sources fail or if all FLCC branches are off, thetoe brakes are totally inoperative.
Channels 1 and 2 use separate redundant circuit
elements for controlling the brakes and operate thesame except that when CHAN 1 is selected, both dc
power sources are switched off when the LG handle
is up. With CHAN 1 selected, the brakes only operatewith the LG handle down; with CHAN 2 selected, thebrakes are operable with the LG handle either up or
down. If the LG handle is stuck in the UP position,
CHAN 2 must be selected to achieve braking.
SPIN DOWN BRAKING SYSTEM

1F-16X-1-0026X

Figure 139.

The electrical power sources are grouped to provide
two redundant channels. Channel 1 uses FLCC
branches A and C and CHAN 1 dc power. Channel 2
uses FLCC branches B and D and CHAN 2 dc power.
If one FLCC branch fails, one toe brake in either
CHAN 1 or CHAN 2 is inoperative.
An inoperable FLCC branch may illuminate theFLCS FAULT caution light and generate a BRK PWR
DEGR PFL. FLCS PWR lights on the TEST switch
panel should be used to determine the proper
BRAKES channel switch position.
After engine shutdown (main and standby generators
and FLCS PMG not operating), the brakes remain
powered as long as the MAIN PWR switch is not
moved out of MAIN PWR. If the MAIN PWR switch
is moved to BATT, the FLCS relays open and the
FLCC is no longer powered by the aircraft battery.
Therefore, the toe brakes are inoperative.

The spin down braking system provides hydraulicbrake pressure to stop MLG wheel spin during LG
retraction. The hydraulic pressure is relieved whenthe LG is up and locked.
BRAKES CHANNEL SWITCH C

DFThe BRAKES channel switch, located on the LG
control panel, has positions of CHAN 1 and CHAN 2
and allows wheel brake system switching. CHAN 1 isthe normal position.
PARKING BRAKE C

DFThe parking brake is activated by the ANTISKID
switch located on the LG control panel, and supplies
full, unmetered pressure to three of the six pistons ineach brake. The parking brake holds the aircraft
stationary without the use of toe brakes. It can also be
used for emergency braking if the toe brakes areinoperative. The parking brake is powered by battery
bus No. 2 and system B hydraulics or one brake/JFS
accumulator (the brake/JFS accumulator which is
not used for START 1).

1111

T.O. GR1F16CJ1

Wheel Brake Schematic (Typical)
THROTTLE
POSITIONSWITCHNEAR IDLE>1 INCH ABOVE
IDLE

28 VDCBATTERY
BUS
NO. 2

ANTISKID
CAUTION
LIGHT
RELAY
ANTI
SKID

MLG WOWGROUND
AIR

PARK
VALVE

CHANNEL 2
28 VDCBATTERY
BUS
NO. 1

CHANNEL 1

BRAKE CONTROL/
ANTI-SKID ASSYCHANNEL SELECTOR

BRAKE
CHAN 1CA

CHANNEL 1
ORCHANNEL 2OUTPUTSHUTTLE
VALVE

BRAKE
CHAN 2D

B

FLCS
INVERTERSPREBRAKE
RESTRICTOR
VALVE

WHEEL
LG UP
PRESSURE SPEEDSENSORL

R

TOE
BRAKES

26 VAC 800 H z
4 INDEPENDENT EXCITATION POWER
SOURCES FROM FLCS INVERTERS

GR1F-16CJ-1-1042-1X37

Figure 140.(Sheet 1)

1112

T.O. GR1F16CJ1

HYD
RETURNHYD SYS B PRESSURE

LEGEND:
ELECTRICALMECHANICAL

ACCUMULATORHYD POWER
SUPPLYHAND
PUMP

HYD RETURN
BRAKE HYDCIRCUIT

ACCUMULATORTO JFSSUPPLY HYDRAULIC CIRCUITSPARKING BRAKECIRCUITCHECK VALVESHUTOFF VALVESHUTOFF VALVERIGHTMETERING
VALVE

LEFT
METERING
VALVE

SHUTTLEVALVESHUTTLE
VALVE

LEFT
BRAKE

RIGHT
BRAKE

WHEEL
SPEED
SENSOR

GR1F-16CJ-1-1042-2X37

Figure 140.(Sheet 2)

1113

T.O. GR1F16CJ1

Wheel Brake Schematic (Typical) LESSTHROTTLE
POSITIONSWITCHNEAR IDLE

MLG WOWGROUND>1 INCH ABOVE
IDLE

AIR

ANTISKIDCONTROL
BOX

28 VDCBATTERY
BUS
NO. 2

ANTISKID
CAUTION
LIGHT
RELAY
ANTI
SKID

PARK
VALVE

CHANNEL 2
28 VDCBATTERY
BUS
NO. 1

CHANNEL 1
BRAKE CONTROLBOX
CHANNEL SELECTOR

BRAKE
CHAN 1CA

CHANNEL 1
ORCHANNEL 2OUTPUTSHUTTLE
VALVE

BRAKE
CHAN 2D

B

FLCS
INVERTERSPREBRAKE
RESTRICTOR
VALVE

WHEEL
LG UP
PRESSURE SPEEDSENSORL

R

TOE
BRAKES

26 VAC 800 H z
4 INDEPENDENT EXCITATION POWER
SOURCES FROM FLCS INVERTERS

GR1F-16CJ-1-0042-1A37

Figure 140.(Sheet 3)

1114

T.O. GR1F16CJ1

HYD
RETURNHYD SYS B PRESSURE

ACCUMULATOR

HAND
PUMP

ACCUMULATORTO JFSSUPPLY HYDRAULIC CIRCUITSSHUTOFF VALVESHUTOFF VALVERIGHTMETERING
VALVE

LEFT
METERING
VALVE

SHUTTLEVALVESHUTTLEVALVE

LEGEND:
ELECTRICALMECHANICAL

LEFT
BRAKE

RIGHT
BRAKE

WHEEL
SPEED
SENSORHYD POWER
SUPPLYHYD RETURN
BRAKE HYDCIRCUIT
PARKING BRAKECIRCUITCHECK VALVE

GR1F-16CJ-1-0042-2X37

Figure 140.(Sheet 4)

1115

T.O. GR1F16CJ1ANTISKID SYSTEM
The antiskid system is available in either brakechannel anytime the toe brakes are powered.
b2t The antiskid system will deliver a corresponding

deceleration rate to a given pedal deflection. The
deceleration skid control will dampen brake pedal
inputs to the brakes resulting in a smoother, more
efficient stop than with previous antiskid systems. To
optimize braking performance and reduce wear on
aircraft brakes and tires, smoothly apply brakes in asingle application.
Functions are:
D Touchdown skid control - Prevents brake applica

tion prior to wheel spinup even if brake pedals are
fully depressed.

D b2t Deceleration skid control-Active when either

brake pedal deflection is less than 85 percent of
maximum and runway surface can provide therequested deceleration.

D b2t Maximum performance skid control-Active

when both brake pedal deflections are equal to or
greater than 85 percent or runway surface cannot
provide requested deceleration.

D b2t Antiskid failure detection-Detects a failure

affecting braking or in a system component.

D LESSb2t Proportional skid control-Prevents
skidding due to overbraking at 5 knots groundspeed
or greater.

D LESS b2t Locked wheel skid control-Backs up the

proportional skid control and operates at 20 knots
groundspeed and greater.

D LESS b2t Antiskid failure detection-Detects anantiskid system malfunction.

b2t If a failure affecting braking performance is
detected while the aircraft is moving above 5 knots
groundspeed, the ANTI SKID caution light illumi
nates. In most cases this represents the loss of a wheel
speed sensor signal, and the system switches to analternate braking mode. In this mode, if differential
braking is applied (15 percent or greater difference
between pedals), both brakes alternate between
pedal pressure as metered and no pressure. Braking
effectiveness is reduced by 50 percent or greater. If
brake pedals are within 15 percent, the system uses
the information from the remaining good wheel speed
sensor and stopping distance is increased byapproximately 25 percent on both wet and dry
runways.

1116

b2t The alternate mode continues until the BRAKES

channel switch is switched to CHAN 2 and the
ANTISKID switch is placed to OFF. The ANTI SKID
caution light remains on and braking is manual. The
brakes then can be locked by applying too much pedal
pressure, which may result in blown tires.
b2t The antiskid system incorporates a hydroplaning

protection function which prevents brake application
until the wheels have spun up, even if WOW has
occurred before spinup.
b2t Full antiskid function becomes active at 12 knots

groundspeed when accelerating and is available tobelow 5 knots when decelerating. Maximum braking
below 12 knots groundspeed may result in tire flatspotting.
LESS b2t If a failure is detected, the ANTI SKID
caution light illuminates and the brake system
automatically switches to pulsating pressure(constant frequency pulsating onoff pressure). In
this mode, braking effectiveness is reduced approxi
mately 50 percent; however, in most cases, braking
effectiveness is as good as can be obtained with
ANTISKID switch in OFF while avoiding wheel
lockup and its associated risk of control difficulty.
Short field landing distances are increased approxi
mately 60 percent for dry runway and 25 percent for
wet runway from those normally computed. The
amount of pulsating braking is dependent on the toepressure applied. Pulsating braking continues until
the ANTISKID switch is placed to OFF. At that time,
the ANTI SKID caution light remains on, and thebrake system reverts to manual control. The brakes
then can be locked by applying too much pedal
pressure, which may result in blown tires.
LESS b2t The antiskid system does not provide skid or
locked wheel protection if MLG wheels are notspinning due to hydroplaning. If WOW occurs prior to
spinup of at least one MLG wheel, wheel brakes
become operative without antiskid protection.
LESS b2t Antiskid protection and failure detection areintentionally diminished as speed decreases below 20
knots to allow a complete stop without continuous
brake release or an ANTI SKID caution light.
Consequently, maintaining maximum toe pressure
while the aircraft comes to a complete stop may cause
wheel lockup during the last 510 feet before stoppingwith resultant tire flat spotting. A dragging brake notcorrectly releasing in response to antiskid signals
may also cause wheel lockup without an illuminated
ANTI SKID caution light at speeds below 20 knots.
Avoid maximum braking at speeds below 20 knots
whenever possible to prevent tire flat spotting andpossible blowout.

T.O. GR1F16CJ1
ANTISKID Switch C

DF

The ANTISKID switch, located on the LG control
panel, is not leverlocked in the ANTISKID positionand can be bumped to OFF.
Functions are:
D PARKING BRAKE - Full unmetered brake pressure

is applied with the throttle in the OFF to IDLE range
and WOW. Advancing the throttle more than 1 inch
beyond IDLE automatically returns the switch to
ANTISKID which releases the parking brake.

D ANTISKID - Antiskid protection is available.
D b2t OFF-Parking brake feature is deactivated andantiskid functions are as follows:
S WithBRAKES channel switch in CHAN 1Touchdown skid control is not available, but
deceleration and maximum performance skid
control remain active.

S With BRAKES channel switch in CHAN 2-Allantiskid functions are deactivated.

D LESSb2t OFF-Antiskid and parking brakefeatures are deactivated.

ANTI SKID Caution Light
b2t The ANTI SKID caution light, located on thecaution light panel, illuminates at groundspeeds
above 5 knots when a malfunction affecting braking
performance is detected. If a system malfunction notaffecting braking performance (e.g. loss of redundancy) is detected, the light illuminates whengroundspeed is below 5 knots. The caution light is notlatched and will extinguish above 5 knots if a failurethat does not affect braking performance is present.
b2t The ANTI SKID caution light illuminates whenpower is applied to the brake control/antiskid
assembly and goes off when powerup BIT has beensuccessfully completed (approximately 1/2 second
later). This brief illumination of the ANTI SKID
caution light may be observed when power is first
applied or after the LG handle is placed down with the
BRAKES channel switch in CHAN 1.
LESS b2t The ANTI SKID caution light, located on thecaution light panel, illuminates when a malfunction
occurs with the ANTISKID switch in ANTISKID.
The ANTI SKID caution light illuminates when the
LG handle is down and the switch is in OFF.

SPEEDBRAKE SYSTEM
The speedbrake system consists of two pairs ofclamshell surfaces located on each side of the engine
nozzle and inboard of the horizontal tail and is
powered by hydraulic system A. The speedbrakes
open to 60 degrees with the right MLG not down andlocked. With the right MLG down and locked,
speedbrake opening is limited to 43 degrees toprevent the lower surfaces from striking the runway
during landing. This limit can be overridden byholding the SPD BRK switch in the open (aft)position. When the NLG strut compresses on landing,the speedbrakes can be fully opened and remain fully
open without holding the SPD BRK switch.
SPD BRK Switch
C DF The SPD BRK switch, located on the throttle,
is a thumbactivated, threeposition slide switch. The
open (aft) position is springloaded to off (center) andallows the speedbrakes to be incrementally opened.
The closed (forward) position has a detent, allowing
a single motion to close the speedbrakes. To prevent
possible creeping, the switch should be left in theclosed position. DR For SPD BRK switch differences,refer to F16D AIRCRAFT, this section.
D The speedbrake switches are connected in paralleland function so that either can override the other byholding in the open position. If one switch is in theclosed position, the speedbrakes close when the otheris released from the open position.

SPEED BRAKE Position Indicator
A threeposition SPEED BRAKE indicator is locatedon the LG control panel.
Positions are:
D CLOSED - Both speedbrakes closed.
D Speedbrake symbol - Speedbrakes not closed.
D Diagonals - Electrical power removed from theindicator. Diagonals also appear momentarily
during switching.

DRAG CHUTE SYSTEM
A drag chute is provided to minimize stopping
distance. Drag chute deployment is obtained when
hydraulic system B pressure is routed to the dragchute actuator by placing the DRAG CHUTE switch
to DEPLOY. Drag chute accumulator pressure isavailable in case of hydraulic system B failure.
1117

T.O. GR1F16CJ1Extension of the actuator closes a set of jaws onto theparachute Dring and pulls the ripcord that releases
a springloaded pilot chute. The pilot chute functions
to pull the main canopy from deployment bag locatedin an aerodynamic fairing below the rudder.
Deployment below 90 knots may result in improper
deployment and damage to the drag chute.
The drag chute system has safety provisions for
accidental deployments, both commanded anduncommanded:
D Above 190 knots , the mechanical fuse section on theover the cable. D Either HOOK switch may be used
to extend the hook. Both switches must be positioned
to UP to raise the hook.
HOOK Caution Light
The HOOK caution light, located on the caution light
panel, illuminates anytime the hook is not up andlocked.

WING FLAP SYSTEM

Dring fails and releases the chute.

D A deployed drag chute (or some residue thereof)

resulting from placing the DRAG CHUTE switch to
DEPLOY can be released by moving the switch
to C DF NORM/REL or DR REL at any airspeed.

D If the drag chute is deployed uncommanded (i.e.,

ripcord failure) and airspeed is above 60 knots, theDring pulls out of the jaw mechanism.

DRAG CHUTE Switch

LEADING EDGE FLAPS (LEF'S)
The LEF's consist of a spanwise flap on each wingleading edge controlled as a function of mach number,
AOA, and altitude by command signals from the
FLCC.
An asymmetry sensing and braking mechanism
prevents LEF asymmetry. If an asymmetry is sensed,
the LEF's lock, FLCS LEF LOCK PFL is displayed,
and the FLCS warning light illuminates.

Refer to figure 141. C DF The DRAG CHUTE switch,
located on the MISC panel, is a twoposition guarded
switch used to deploy and release the drag chute. The
switch is powered by battery bus No. 1. DR For DRAG
CHUTE switch differences, refer to F16D AIR
CRAFT, this section.

The LEF's are automatically programmed when the
LE FLAPS switch is in AUTO.

ARRESTMENT SYSTEM

D When the throttle is at IDLE and MLG wheel speed

The hook is electrically controlled and pneumaticallyoperated. Pneumatic pressure is supplied by the
LG/hook emergency pneumatic bottle which contains
sufficient pressure to lower the LG and hook.

D LEF asymmetry brakes are locked.

When extended, pneumatic pressure holds the hookon the runway. When subsequently retracted, thehook rises enough to allow the cable to drop off thehook or to be disengaged. The hook is springloaded
partially up to allow taxiing over a cable. The hookmust be raised manually to reset it to the stowed
position.
HOOK Switch
The HOOK switch , located on the LG control panel,is leverlocked in the UP or DN position. Positioning
the switch to DN causes the hook to extend.
Returning the switch to UP partially retracts thehook, allowing for cable disengagement and for taxi

1118

Exceptions are:
D When weight is on both MLG (the LEF's are 2

degrees up).

is greater than 60 knots groundspeed (the LEF's are
2 degrees up).

D When the FLCS is operating on standby gains.

Refer to STANDBY GAINS, this section.

LE FLAPS Switch
The LE FLAPS switch is covered as a part of the FLT
CONTROL panel.
FLCS LEF LOCK PFL
The FLCS LEF LOCK PFL is activated bymalfunctions in the flap drive unit or flap commands.
The FLCS LEF LOCK PFL is also activated if the
LEF's are manually locked or the asymmetry brakes
are activated. The FLCS warning light illuminates
when the FLCS LEF LOCK PFL occurs.

T.O. GR1F16CJ1Drag Chute Controls (Typical)
C

DF

DR

1

2
NOTE:
DR For DRAG CHUTE switch,refer to F-16D AIRCRAFT, thissection.

1. C DF DRAG CHUTE Switch
2. DR DRAG CHUTE Switch

GR1F-16CJ10043X37

SWITCHPOSITIONNORM

FUNCTION
Chute stowed
Attachment jaws at midrange

DEPLOY

Hydraulic actuator extends
Ripcord pulled to release springloaded pilot chuteAttachment jaws close on Dring

REL

Hydraulic actuator retracts for 2 seconds, releasing the Dring from the attachment jaws
and then returns to midrange

Figure 141.

1119

T.O. GR1F16CJ1TRAILING EDGE FLAPS (TEF'S) (FLAPERONS)
The flaperons are located on the wing trailing edge
and function as ailerons and TEF's. The flaperons
have a maximum command deflection of 20 degrees
down and 23 degrees up. When acting as flaps, thedeflection is downward; when acting as ailerons, thedeflection is up or down, as commanded. Bothfunctions are operable whenever the FLCS ispowered. The TEF's are controlled as a function of the
LG handle position, the ALT FLAPS switch, airspeed,
and mach number. Positioning the LG handle to DN
or the ALT FLAPS switch to EXTEND causes the
TEF's to deflect downward. At all airspeeds below 240
knots, the TEF's position is 20 degrees down. Above
240 knots, the TEF's reduce deflection as a functionof airspeed until nearly/fully retracted at 370 knots.
ALT FLAPS Switch
The ALT FLAPS switch is located C DF on the FLCP,DR on the LG control panel. With the switch in
NORM, the TEF's are controlled by the LG handle
and airspeed. Placing the switch to EXTEND lowers
the TEF's only, depending on airspeed. The ALT
FLAPS switch does not affect the operation of the
LEF's unless the FLCS is operating on standby gains.
Refer to STANDBY GAINS, this section.

FLIGHT CONTROL SYSTEM (FLCS)
Refer to figure 142 for FLCS functional schematicand figure 143 for FLCS pitch, roll, and yaw
schematic. The FLCS is a digital fourchannel,
flybywire system which hydraulically positions
control surfaces. Electrical signals are generated
through a stick, rudder pedals, and a MANUAL
TRIM panel. A main component of the FLCS is theflight control computer (FLCC). Redundancy isprovided in electronic branches, hydraulic systems,
power supplies, and sensor systems. A FLTCONTROL panel (FLCP) provides a BIT RUN FAIL
light and controls.
Command signals to the FLCC are initiated byapplying force to the stick and rudder pedals. Thesesignals are processed by the FLCC along with signals
from the AOA sensors, air data system, flight control
rate gyros, accelerometers, and INS. The processedsignals are transmitted to the ISA's of the horizontal
tails, flaperons, and rudder which are positioned togive the commanded response.
Pitch motion is controlled by symmetrical movementof the horizontal tails. Roll motion is controlled bydifferential movement of the flaperons and horizontal
tails. Yaw motion is controlled by the rudder. Roll
1120

coordination is provided by an ARI. The ARI function
is not available whenever MLG wheel speed exceeds
60 knots or if AOA exceeds 35 degrees. After takeoff,
ARI is activated within 2 seconds after the LG handle
is raised (spin down braking system). If the LG
handle remains down, 1020 seconds are required for
the MLG wheels to spin down and activate ARI.
Digital Backup (DBU)
DBU provides a software backup in the event ofsoftware problems in the primary program. The DBU
is a reduced set of control laws which automaticallyengages when software problems in the FLCC force amajority of the branches into a failed state. DBU can
only be disengaged by use of the DIGITAL BACKUP
switch.
DBU operation does not significantly impact aircraft
handling qualities during normal cruise operation or
landing.
During DBU operation:
D The AOS feedback function is inoperative.
D Autopilot, terrain following, and stick steering areinoperative.

D Gun compensation is not provided.
D With the LG handle in DN, the TEF's are positioned

to 20 degrees down. With the LG handle in UP and
the ALT FLAPS in EXTEND, the TEF's are
positioned to 20 degrees down if airspeed is lessthan 290 knots.

D There is no roll rate input to the AOA limiter.

Maximum roll rate command is a constant 167
degrees per second.

D STORES CONFIG switch is inoperative.
D Stick commands are essentially CAT III limited.

Rudder pedal commands are essentially CAT I
limited.

D Dual air data failures are not recognized in DBU.

Standby gains are not engaged. If the LG handle isin UP, midvalue air data is always selected. If the
LG handle is in DN, gains are fixed at normal
landing values.

D Pitch trim centering at wheel spinup is inopera

tive.

D AOA indications do not set to zero and random AOA

indications are possible in gusty wind conditions
with NLG WOW.

T.O. GR1F16CJ1
D D The STICK CONTROL switch and the overridefeature of the paddle switch are inoperative. If MPO
is activated, it cannot be overridden in the othercockpit.

D LEF scheduling is simplified and optimized for a

cruise condition at approximately 20,000 feet MSL.

D LEF are not commanded to 2 degrees up when MLG

wheel speed is greater than 60 knots and thethrottle is IDLE or if MLG WOW.

D DBU does not support communication on the MUXbus, so failures while in DBU are not reported.

FLCS LIMITERS
Refer to figure 144 for limiter values. FLCS limiters
are provided in all three axes to help prevent
departures/spins.
AOA/G Limiter
Refer to figure 145. In cruise gains, the AOA/g limiter
reduces the positive g available as a function of AOA.
The negative g available is a function of airspeed.
Below 15 degrees AOA, the maximum positive gavailable is +9g. As AOA increases, the maximum
allowable positive g decreases. The positive g limit
and maximum AOA depend on the position of the
STORES CONFIG switch. In CAT I, positive gdecreases to a value of 1g at 25 degrees AOA.
Maximum commanded AOA is approximately 25.5
degrees. In CAT III, maximum AOA varies fromapproximately 1618 degrees as a function of GW andg.
The negative g available above approximately 250
knots is -3g. Below 250 knots, the available negative
g varies between 3g and zero g as a function of
airspeed, altitude, and AOA.
In takeoff and landing gains, the STORES CONFIG
switch has no effect on limiting or gains. Maximum
positive g is a function of airspeed and AOA. The
negative g command limit is not a function of
airspeed. It is a fixed limit. The maximum AOA for 1g
is approximately 21 degrees.
In inverted or upright departures, the AOA/g limiter
will override stick pitch commands if the MPO is notengaged. The MPO can always override the negative
g function of the limiter. It can also override the AOA
function of the limiter when the AOA exceeds 35
degrees. Refer to MPO, this section.

Roll Rate LimiterIn cruise gains, the roll rate limiter reduces available
roll rate authority to help prevent roll coupled
departures. This authority is reduced as airspeed
decreases, AOA increases, or trailing edge down
horizontal tail deflection increases. Roll authority isfurther reduced for large total rudder commands. In
takeoff and landing gains, roll rate limiting isavailable but is a fixed value independent of AOA,
airspeed, or horizontal tail position.
Rudder Authority LimiterIn cruise gains, the rudder authority limiter reduces
the pedal commanded rudder deflection as a functionof AOA, roll rate, and STORES CONFIG switch
position for departure protection. However, ARI
authority, stability augmentation, and trim authority
are not reduced. In takeoff and landing gains,
category I rudder authority limiting is provided.
Yaw Rate Limiter
When AOA exceeds 35 degrees, the yaw rate limiter
overrides pilot roll and rudder commands andprovides flaperon with and rudder against the yaw
rate until AOA is below 32 degrees to enhance spin
resistance. The yaw rate limiter provides noprotection against yaw departures in the normal
flying range (-5 to 25 degrees AOA).
When AOA decreases below -5 degrees and airspeed
is less than 170 knots, the yaw rate limiter engages
but does not affect pilot roll and rudder commands.
Pilot roll and rudder commands are inhibited during
inverted departures only when the MPO is engaged.
The yaw rate limiter provides rudder against the yaw
rate until AOA is above -5 degrees to enhance spin
resistance.
Automatic yaw rate limiting to enhance spin
resistance is independent of the angleofsideslip
feedback function; thus, limiting is available even if
the FLCS AOS FAIL PFL is present.
FLCS GAINS
During normal operation, the FLCS receives inputs(gains) from the ADC and provides relatively
constant aircraft response for a given stick input,
regardless of altitude or airspeed. This response
varies slightly depending on configuration. In theevent of a dual air data failure, the FLCS switches tostandby (fixed) gains.

1121

T.O. GR1F16CJ1FLCS Functional Schematic (Typical)
COMMANDS
Cockpit
BIT

TRIM INPUTS
Side Panel
Stick
MANUAL PITCH
Override

FLCS POWER INPUTSCONTROL INPUTS
Stick
Rudder Pedals
RATES
Roll
Pitch
Yaw

FLIGHT CONTROL PANELSWITCHESBIT INDICATIONMANUAL TRIM PANEL
TRIM CONTROLSTRIM/AP DISCPITCH TRIM CENTERINGMANUAL PITCH OVERRIDESWITCH PANELDC POWER FLCS
CONVERTER/REGULATORSQUADRUPLE FORCESENSORS

3-AXIS COMMANDSTEF COMMANDQUADRUPLE RATESENSORSACCELERATION
NormalLateralQUADRUPLEACCELEROMETERS

AIRFLOWAOA

DUAL AIRFLOW SENSORSAIR DATA
Total Pressure
Static Pressure
AOA

AIRCRAFT MANUAL TRIMPNEUMATIC SENSORASSEMBLY (PSA)

AIR DATA
Total Pressure
Static Pressure
Total Temperature

CENTRAL AIR DATACOMPUTER (CADC)

AIR DATA
Static Pressure

SIDESLIP DIFFERENTIALPRESSURE SENSOR

AOA

GR1F16CJ100441X37

Figure 142.(Sheet 1)

1122

T.O. GR1F16CJ1WARNING and CAUTION
LIGHTSRIGHT HORIZONTAL TAIL
INTEGRATEDSERVOACTUATOR (ISA)
FLIGHT CONTROL COMPUTERSTABILITY/COMMAND
AUGMENTATIONQUADRUPLE ELECTRONICS/
PROCESSORSCONTROL DYNAMICSGAIN SCHEDULINGQUADRUPLE INPUT/OUTPUT SELECTORS
800 HZ SENSOR POWER
SUPPLYAILERON-RUDDER
INTERCONNECT
SIDESLIP FEEDBACK
STRUCTURAL FILTERSAOA LIMITINGYAW RATE LIMITINGHIGH AOA ROLL RATE
LIMITING
DIGITAL BACK-UP
AUTOPILOTAUTO TF

LEFT HORIZONTAL TAIL
INTEGRATEDSERVOACTUATOR (ISA)

RIGHT FLAPERONINTEGRATEDSERVOACTUATOR (ISA)

LEFT FLAPERONINTEGRATEDSERVOACTUATOR (ISA)RUDDERINTEGRATEDSERVOACTUATOR (ISA)SENSOR POWER AND TESTAMUXPX II FCC, PX III MMCMFL'S, PFL'SINS PITCH, ROLL,YAW ANGLES;INERTIAL VELOCITIESCADC-ALTITUDE
NVP-TF COMMANDSDUAL LEF COMMANDSERVOSLEF ACTUATIONSYSTEMDMUX
MFD FLCS PAGEPFL'SMEMORY MODE

LEGEND:

ELECTRICALMECHANICAL

GR1F-16CJ-1-0044-2A37

Figure 142.(Sheet 2)

1123

T.O. GR1F16CJ1FLCS Pitch, Roll & Yaw Schematic (Typical)
PITCH
PITCH TRIMMECHANICAL LIMITHORIZONTAL TAIL STICK FORCEFLIGHT
CONTROLCOMPUTERAOAPITCH RATEISARIGHT HORIZONTAL
TAIL

ISA

LEFT HORIZONTAL
TAIL

ISA

RIGHT
FLAPERON

ISA

LEFT
FLAPERONISARUDDERNORMAL ACCELERATIONIMPACT PRESSURE

ROLL
ROLL TRIMMECHANICAL LIMITAILERON STICK FORCEROLL RATEYAWYAW TRIMMECHANICAL LIMITRUDDER PEDAL FORCEYAW RATELATERAL ACCELERATIONGUN COMPENSATIONNOTES:FLCS gains scheduled by air
data inputs.

GUN FIRE

Sideslip angle and rate
calculated in the FLCC
based on INS inputs.

INS
ANGLES AND VELOCITIESLEGEND:
ELECTRICALMECHANICAL
1F-16X-1-4009X

Figure 143.

1124

T.O. GR1F16CJ1FLCS Limiter Functions
PITCH AXIS
Maximum AOA=25°

CAT
I

g command system until 15° AOA

g/AOA command system above
15° AOA

Maximum AOA=16°18°
pending on GW)

CAT
III

(de

g command system until 7° AOA
at 100 knots to 15° AOA at 420
knots and above

ROLL AXISYAW AXIS

Maximum roll rate commanddecreases with:

Maximum deflection (pedal command) reduced for:D AOA above 15°

D AOA>14° (zero roll rate)

D Airspeed less than 250 knots

D Roll rate>20°/sec

D Horizontal tail deflection morethan 5° trailing edge down

NOTE:
Zero rudder authority available at
26° AOA

D Total rudder command (frompilot and FLCS) exceeding 20°
D Combination of horizontal tailgreater than 15° trailing edge
down and AOA above 22°
Maximum roll rate command
reduced by approximately 40 per
cent of CAT I authority. Additional
decreases as function of AOA, airspeed, horizontal tail position,and total rudder command

NOTES

2. +9g available until 15° AOA.
Maximum g decreases as au c o o
O a
da
speed
functionof AOAandairspeed

D AOA>3° (zero roll rate)
D Roll rate>20°/sec
NOTE:
Zero rudder authority available at
15° AOA

g/AOA command system above
these values
1. In takeoff/landing gains, theFLCS operates as a pitch ratecommand system until 10°
AOA and a pitch rate/AOA
command systems stem above
abo e 10°
AOA

Maximum deflection (pedal command) reduced for:1. In takeoff/landing gains, maxi
mum roll rate is fixed at
approximately onehalf the
maximum roll rate available incr ise gains,
cruisegains regardless of
AOA, airspeed, or horizontal
tail deflection
2. Above 35° AOA, the yaw rate
limiter cuts out stick roll com
mands
d and
d provides
id rollll axisi
antispin control inputs

1. Above 35° AOA, the yaw rate
limiter provides yaw axis
antispin control inputs
2. Below 5°5 AOA and less than
170 knots, the yaw rate limiter
provides antispin rudder in
puts; pilot roll and rudder commands are cut out only wheng g
MPO is engaged
3. Maximum deflection (30°)
always available thru ARI andstability augmentation

Figure 144.

1125

T.O. GR1F16CJ1

AOA/G Limiter Function (Cruise Gains)

10

9
CAT I SELECTEDNORMAL LOAD FACTOR

G

8

7

6

CAT III SELECTEDGW = 25,000 POUNDS

5

CAT III SELECTEDGW = 30,000 POUNDS

4

3

CAT III SELECTEDGW = 35,000 POUNDS

2

1
15

16

17

18

19

20
AOA

21

22

23

24

25

26

DEGREES

1F-16X-1-0010X

Figure 145.

1126

T.O. GR1F16CJ1
Cruise Gains
The FLCS is in cruise gains with the LG handle in UP,
the ALT FLAPS switch either in NORM or in
EXTEND above 400 knots, and the AIR REFUEL
switch either in CLOSE or in OPEN above 400 knots.
At low AOA, the pitch axis of the FLCS is a g
command system. As AOA increases, the FLCS
switches to a blended g and AOA system to provide a
warning of high AOA/low airspeed. Roll rate limiting
is available and maximum roll rate decreases as afunction of low airspeed, high AOA, and horizontal
tail position.
Takeoff and Landing Gains
The FLCS is in takeoff and landing gains with the LG
handle in DN, the ALT FLAPS switch in EXTEND(below 400 knots), or the AIR REFUEL switch in
OPEN (below 400 knots). In takeoff and landing
gains, the FLCS pitch axis operates as a pitch ratecommand system until 10 degrees AOA and a blended
pitch rate and AOA command system above 10
degrees AOA. Roll rate limiting is available but is a
fixed value independent of AOA, airspeed, or
horizontal tail position.

ANGLEOFSIDESLIP (AOS) FEEDBACK FUNC
TION
The angleofsideslip feedback function provides
improved departure prevention by using AOS andAOS rate feedback to position the rudder. AOS andAOS rate are calculated in the FLCC using INS data.
The calculated AOS is also monitored in the FLCC bycomparing it to an AOS derived from differential
pressure sensor signals. This monitoring detects oneof two possible failures:
D AOS derived from INS is erroneous.
D AOS derived from differential pressure sensor iserroneous.

Either failure deactivates the AOS feedback function
and activates the FLCS FAULT caution light and theFLCS AOS FAIL PFL.
The AOS feedback function is active when all of thefollowing conditions are met:
D Airspeed is less than 350 knots.
D AOA is greater than 10 degrees.
D AOS exceeds 2 degrees.

Standby Gains

D AOS monitoring has not detected a failure.

In standby gains, control response is tailored for a
fixed altitude (sea level, standard day) and airspeed(LG handle in UP, approximately 600 knots; LG
handle in DN, approximately 230 knots). The FLCS
warning light and FLCS FAULT caution lightilluminate.

D DBU is not engaged.

When operating on standby gains, the LEF's are atzero degrees with the LG handle in UP and the ALT
FLAPS switch in NORM. The LEF's deflect 15
degrees down with the LG handle in DN or the ALT
FLAPS switch in EXTEND. The operation of the
TEF's is not affected in standby gains.

GUN COMPENSATION

A standby gains condition can be reset in flight, backto the first failure condition, by using the FLCSRESET switch. The original air data system failure islatched upon occurrence of the second failure anddoes not reset. If reset is successful, the FLCS
warning light goes off.
FLCS DATA RECORDER C

DFThe FLCS data recorder is attached to the ejection
seat and departs the aircraft on ejection. It retains thesame information as the FLCC including FLCS
failure data, airspeed, altitude, true heading andelapsed time from takeoff.

D MPO switch is in NORM.
D The FLCS is in cruise gains.
D Terrain following is not engaged.

The FLCS automatically compensates for the off
center gun and the aerodynamic effects of gun gas
emissions during firing by moving the flaperons andrudder. Gun compensation is optimized for 0.70.9
mach range; therefore, all excursions may not be
eliminated. For example, gunfiring at low mach may
result in nose left excursions while nose right
excursions are likely at higher mach. Failure
monitoring of gun compensation circuits is notprovided and there are no caution light indications forincorrect compensation.
FLIGHT CONTROL SYSTEM (FLCS) CONTROLS
StickRefer to figure 146. The stick is a forcesensing unit
which contains transducers in both pitch and rollaxes, moves approximately 1/4 inch in both axes, andis rotated slightly cw.
Change 11127

T.O. GR1F16CJ1

Stick (Typical)
2

1

NWS
A/R DISCMSL STEP

3

9CAMERA/
GUN

4
WPN
REL

8

TRIM

5
7

6

1. NWS A/R DISC MSL STEP Button
2. TRIM Button (4-Way, Momentary)
3. Display Management Switch(4-Way, Momentary)
4. Target Management Switch(4-Way, Momentary)
5. Countermeasures Management
Switch (4-Way, Momentary)
6. Expand/FOV Button
7. Paddle Switch
8. CAMERA/GUN Trigger (2-Position)
9. WPN REL Button

GR1F-16CJ-1-0047X37

CONTROLPOSITION

1. NWS A/R DISC MSLSTEP Button(NWS)

Depress (on ground)

Activates NWS

Depress (2nd time)

Deactivates NWS(A/R DISC)

Depress (in flight)

Disconnects boom latching. AIR REFUEL
switch must be in OPEN position(MSL STEP)

Depress (in flight)

Activates missile step function. Refer toT.O. GR1F16CJ3411

Fwd
Aft
Left
Right

Trims nosedown
Trims noseup
Trims left wing down
Trims right wing down

2. TRIM Button (NOSE DOWN)(4way, momentary) (NOSE UP)
(LWD)
(RWD)

Figure 146.(Sheet 1)

1128

FUNCTION

T.O. GR1F16CJ1

Stick (Typical)
CONTROL
3.

Display Management Switch(4way,(4
way momentary)

POSITIONUp
Down

FUNCTION
Refer to T.O. GR1F16CJ3411 for a de
tailed description of switch functions

Left
Right
4.

Target Management Switch(4way,(4
way momentary)

Up
Down
Left
Right

5.

Countermeasures Manage
ment Switch (4(4way,way momen
momen
tary)

Fwd
Aft
Left
Right

6.

Expand/FOV Button

Depress

Successive depressions sequence throughthe available fieldofview (FOV) selections
for the sensor/system mode being displayedon the DOI

7.

Paddle Switch

Depress

Interrupts the autopilot while switch is depressed. Terminates ATF flyup. If in ATF,
reverts to manual TF while depressed.
Resets SWIM monitors when releasedDFor
stick override function, refer toF16D AIRCRAFT, this section

8.

9.

CAMERA/GUN Trigger(2position)

WPN REL Button

Squeeze trigger todetentStarts operation of AVTR/CTVS with AUTO
selected on AVTR power switch and provides consent for laser fire (if selected andarmed)

Squeeze trigger pastdetentFires gun (if selected and armed), AVTR/CTVS operation continues, and consent for
laser fire continues (camera operation continues for 30 seconds after trigger isreleased)

Depress

Signals consent to PX II FCC or SMS,PX III MMC to initiate weapon release andoperates HUD camera for 30 seconds whenin AUTO

Figure 146.(Sheet 2)

1129

T.O. GR1F16CJ1Maximum noseup and nosedown pitch commands are
generated by 25 and 16 pounds of input, respectively.
Roll commands are generated by a maximum of 17
pounds in cruise gains and by 12 pounds in takeoff
and landing gains. When using the switches/buttons
on the stick, inadvertent commands to the FLCS are
possible.

D AOA is 15 degrees or greater with LG handle down

The wristrest and armrest assemblies which may be
used in conjunction with the stick are located on theright side wall aft of the stick.

D When the TFR is operating and airspeed drops

Rudder Pedals
The rudder pedals are forcesensing units containing
transducers. Force on the applicable rudder pedal
produces electrical yaw command signals.
The rudder pedals are also used to generate brakeand NWS signals. Rudder pedal feel is provided bymechanical springs.
MANUAL TRIM Panel
Refer to figure 147. The MANUAL TRIM panel,
located on the left console, contains trim controls andindicators.
MANUAL PITCH Override (MPO) Switch
Refer to figure 148. The MPO switch, located on theleft console, has two positions, NORM and OVRD,
and is springloaded to the NORM position. This
switch is used during a deep stall condition to enable
manual control of the horizontal tails. Positioning
and holding the switch to OVRD overrides thenegative g limiter. If AOA exceeds 35 degrees, theOVRD position overrides the AOA/g limiter andallows pitch commands.
STORES CONFIG Switch

C DF

Refer to figure 149. The STORES CONFIG switch,
located on the LG control panel, has two positions,
CAT I and CAT III. The CAT III position shall beselected when the aircraft is configured with acategory III loading.
AOA limiting is provided. Refer to FLCS LIMITERS,
this section, for a description of categories I and III
AOA limiter.
Low Speed Warning Tone
Refer to figure 150. A low speed warning tone(steady) sounds in the headset when one of thefollowing conditions exists:

1130

or ALT FLAPS switch in EXTEND.

D Combined airspeed and pitch angle fall on a point

within the tone on area with LG handle up and ALT
FLAPS switch in NORM.
below 300 knots (KCAS or groundspeed).

The low speed warning tone has priority over the LG
warning horn. Depressing the HORN SILENCER
button silences the low speed warning tone.
The low speed warning tone is reactivated only after
the original warning condition is eliminated. The
MAL & IND LTS test button does not test the low
speed warning tone.
FLT CONTROL Panel (FLCP) C

DF

Refer to figure 151. The FLCP, located on the left
console, contains indicators and controls related toflight control functions.
FLCS WARNING, CAUTION, AND INDICATOR
LIGHTS
The instrument panel, caution panel, PFLD, FLCP,ELEC control panel, TEST switch panel, and avionicsystem contain warning, caution, and indicator lights
related to the FLCS. The FLCS is interfaced to theavionic system via the multiplex (MUX) bus toprovide PFL and MFL reporting.
FLCS Warning LightThe FLCS warning light, located on the rightglareshield, illuminates to indicate a dual malfunc
tion in the FLCC electronics, including theprocessors, power supplies, input commands or
sensors, AOA, or air data inputs. The FLCS warning
light also illuminates if the LEF's are locked or BIT
fails. The FLCS warning light remains illuminateduntil FLCS reset action is successful in clearing the
failure. If an active warning fault exists and asubsequent warning level malfunction occurs, theFLCS warning light goes off momentarily to retrigger
HUD WARN and voice warning.
More specific system failure information can beobtained from the PFLD.

T.O. GR1F16CJ1

MANUAL TRIM Panel
1

C DF (Typical)

2

3
1. ROLL TRIM Wheel
2. ROLL TRIM Indicator
3. TRIM/AP DISC Switch
4. PITCH TRIM Indicator
5. PITCH TRIM Wheel
6. YAW TRIM Knob

4

6

5
GR1F-16CJ-1-0048X37

CONTROL
1.

ROLL TRIM Wheel

POSITIONFUNCTION

L WING DN rotation

Trims left wing down

R WING DN rotation

Trims right wing down

2.

ROLL TRIM Indicator

VisualIndicates roll trim

3.

TRIM/AP DISC Switch

NORM

Energizes stick TRIM button. Permits autopi
lot engagement

DISC

Deenergizes stick TRIM button, prevents
autopilot engagement and ATF/autopilot
blending, and deactivates trim motors(manual trim wheels still operative)

4.

PITCH TRIM Indicator

VisualIndicates pitch trim

5.

PITCH TRIM Wheel

NOSE UP rotation

Trims noseup

NOSE DN rotation

Trims nosedown

CCW rotation

Trims nose left

CW rotation

Trims nose right

6.

YAW TRIM Knob

Figure 147.

1131

T.O. GR1F16CJ1Manual Pitch Override Switch (Typical)
C

1

DF

Low Speed Warning
Tone Schedule
NOTES:LG handle in UP and ALT FLAPS switch in NORM.
Dashed lines indicate airspeed tolerances for low
speed warning tone activation.

DR

250
225
200

KIAS

175
150
125

TONE ON

100
75

1. MPO Switch

0
GR1F-16CJ-1-0049X37

Figure 148.

30

45 60 75 90 105 120 135 150
PITCH ATTITUDE DEGREES

1F-16X-1-0022X

Figure150.

STORES CONFIG
Switch C DF
(Typical)DBU ON Warning LightThe DBU ON warning light, located on the rightglareshield, illuminates to indicate the FLCC has
automatically switched to DBU or if the DIGITALBACKUP switch is manually positioned to DIGITALBACKUP.
TO/LDG CONFIG Warning Light
Refer to LANDING GEAR SYSTEM, this section.
FLCS FAULT Caution LightThe FLCS FAULT caution light, located on thecaution light panel, illuminates when the FLCC
reports a caution level PFL item which requires pilot
action. The caution light goes off either when the fault
is acknowledged or when FLCS reset action is
successful in clearing the failure.

1

1. STORES CONFIG Switch

GR1F-16CJ-1-0050X37

Figure 149.

1132

If the FLCS FAULT caution light is illuminated anda subsequent malfunction occurs, the caution lightgoes off momentarily to activate the MASTER
CAUTION light and retrigger voice caution.
More specific system failure information can beobtained from the PFLD.

T.O. GR1F16CJ1FLT CONTROL Panel C DF (Typical)
2

8

3

1

4

10

7

6

MISC Panel

95
1. MANUAL TF FLYUP Switch
2. DIGITAL BACKUP Switch
3. ALT FLAPS Switch
4. RUN FAIL Light
5. BIT Switch

6. FLCS RESET Switch
7. LE FLAPS Switch
8. ADV MODE Switch
9. Autopilot PITCH Switch
10. Autopilot ROLL Switch
GR1F16CJ10052X37

1.

2.

3.

4.

5.

CONTROL/INDICATORPOSITION/INDICATIONMANUAL TF FLYUP
Switch (lever lock)

DISABLEFlyup protection not available in manual TF

ENABLEFlyup protection available in manual TF for
TF or SWIM detected failure

DIGITAL BACKUP Switch(lever lock)

BACKUP

Selects backup software program within the
FLCC

OFF

Normal position

ALT FLAPS Switch(lever lock)

NORM

TEF operation controlled by LG handle

EXTENDTEF's extend regardless of LG handle posi
tions

RUN/FAIL Lights

RUN (green)

Indicates FLCS BIT is running

FAIL (red)Indicates a failure during the FLCS BIT

OFFNormal position

BIT

Commands BIT if weight is on main LG andwheel speed is less than 28 knots
groundspeed

BIT Switch (solenoid
held in BIT andleverlocked to OFF)FUNCTION

Figure 151.(Sheet 1)

1133

T.O. GR1F16CJ1FLT CONTROL Panel C DF (Typical)
6.

7.

8.

9.

10.

CONTROL/INDICATORPOSITION/INDICATIONFLCS RESET Switch(springloaded(springloaded to OFF)OFF

Normal position

RESET

Momentary position which performs servo orelectronic reset of FLCS system failures.
Resets FLCS warning light, CADC, FLCS
FAULT, and MASTER CAUTION lights andclears PFL's if fault is cleared

LE FLAPS Switch(lever lock)

AUTO

LEF's are automatically controlled as a func
tion of mach, altitude, and AOA

LOCKManually locks the LEF's in position and illu
minates the FLCS warning light and FLCS
LEF LOCK PFL

ADV MODE Switch

Autopilot PITCH Switch

Autopilot ROLL Switch

Depressing switch illuminates ATF NOTENGAGED caution light if NVP or FLCS
cannot support ATF
ACTIVE (green)

Depressing switch illuminates ACTIVE lightand selects ATF if TFR is in operate mode

STBY (amber)

Illuminates when ATF/autopilot blended
mode is in use or the paddle switch is depressed in ATF. STBY light indicates that
ATF is ready to take over if the paddle switch
is released or if higher terrain requires it

ALT HOLD

Engages pitch and roll axes of autopilot.
Autopilot maintains constant altitude as
determined by CADC; roll mode is deter
mined by ROLL switch

A/P OFF

Disengages pitch and roll axes of autopilot

ATT HOLD

Engages pitch and roll axes of autopilot.
Autopilot maintains constant pitch attitude
as determined by INS; roll mode is deter
mined by ROLL switch

HDG SEL

Autopilot turns the aircraft to capture andmaintain the heading selected by the head
ing reference marker on the HSI

ATT HOLD

Autopilot maintains roll attitude as deter
mined by the INS

STRG SEL

Autopilot steers aircraft to the selected steer
point

NOTE:
Autopilot roll modes
are functional whenth PITCH switchtheit h iis
out of A/P OFF.

Figure 151.(Sheet 2)

1134

FUNCTION

T.O. GR1F16CJ1
BuiltIn Test (BIT)The BIT switch, located on the FLCP, is a twoposition
switch which is leverlocked in OFF and solenoid heldin BIT. When engaged (green RUN light illuminates),
the BIT runs for approximately 45 seconds. A failure
during the BIT sequence or a BIT interlock failureterminates BIT, returns the switch to OFF,illuminates the red FAIL light and the FLCS warning
light, and sends an FLCS BIT FAIL PFL message.
BIT failures are nonresettable and the red FAIL light
does not go off until a subsequent BIT is successfullycompleted. A BIT detected bus communication failure
results in illumination of the FLCS FAULT caution
light and an FLCS MUX DEGR PFL. An FLCS reset
extinguishes the caution light but the PFL remains.
Successful completion of BIT is indicated by the BIT
switch automatically returning to OFF, the green
RUN light going off, no FLCS BIT FAIL PFL, and a
BIT PASS message on the FLCS page of the MFD.
Pilot Fault List Display (PFLD)The PFLD, located on the C DF right auxiliary
console, DR instrument panel, displays FLCS PFL's.
Refer to WARNING, CAUTION, AND INDICATOR
LIGHTS, this section, for a description of the PFLD.
Refer to PILOT FAULT LIST - FLCS, Section III, fora description of FLCS PFL's.

Rudder trim is initiated from the YAW TRIM knob on
the MANUAL TRIM panel only. For asymmetricconfigurations or roll mistrim, rudder retrimming
may be required as flight conditions change.

AUTOPILOT
The autopilot provides attitude hold, heading select,and steering select in the roll axis and attitude holdand altitude hold in the pitch axis. These modes arecontrolled by PITCH and ROLL switches on the MISC
panel. The TRIM/AP DISC switch on the MANUAL
TRIM panel disengages the autopilot. The paddle
switch on the stick interrupts autopilot operation
while the switch is held depressed.
The PITCH switch is a threeposition switch which issolenoid held in an engaged position and goes to the
A/P OFF position if any of the following conditions
exist.
Conditions are:
D AIR REFUEL switch - OPEN.
D ALT FLAPS switch - EXTEND (below 400 knots).
D A/P FAIL PFL occurs.

TRIM
Pitch trim inputs may be initiated from either the
TRIM button on the stick or the PITCH TRIM wheelon the MANUAL TRIM panel. In cruise gains, pitch
trim inputs are g command signals identical infunction to normal stick pitch commands. In takeoff
and landing gains, the FLCS operates as a pitch ratecommand system until 10 degrees AOA and a pitch
rate and AOA command system above 10 degrees
AOA.
With both main wheel speeds at 60 knots
groundspeed or greater, pitch trim automatically
centers.
Roll trim inputs can be initiated from either the TRIM
button on the stick or the ROLL TRIM wheel on the
MANUAL TRIM panel. For asymmetric configurations or rudder mistrim, roll retrimming may berequired as flight conditions change. Roll trim inputs
also command proportional rudder deflection
through the ARI function.

D AOA greater than 15 degrees.
D DBU - Engaged.
D LG handle - DN.
D Low speed warning tone sounds.
D MPO switch - OVRD.
D STBY GAIN PFL occurs.
D TRIM/AP DISC switch - DISC.

Movement of the PITCH switch out of A/P OFF
engages both the pitch and roll autopilot modes
selected.
The ROLL switch is a threeposition switch whichenables one of the three roll autopilot modes
whenever a pitch autopilot mode is selected.

1135

T.O. GR1F16CJ1
AUTOPILOT OPERATION
The autopilot is fully engaged when the PITCH
switch is not in A/P OFF. Autopilot options are
selected by positioning the PITCH switch (ALT
HOLD, A/P OFF, or ATT HOLD) and the ROLL
switch (HDG SEL, ATT HOLD, or STRG SEL). Sticktrim is inoperative with the autopilot engaged. The
manual trim is operable and may be used while theautopilot is engaged. However, due to the limited
authority of the autopilot, engagement of any modein other than a trimmed flight condition degrades
autopilot performance.
The autopilot loop in the FLCC receives inputs fromthe INS and CADC by means of the AMUX bus. A
lack of data, inaccurate data, or degradation/failureof the AMUX disconnects the autopilot and activates
the FLCS FAULT caution light and the FLCS A/P
FAIL PFL message. The sensor information used by
the autopilot does not include the redundancy of the
FLCS so its use must be closely monitored at low
altitude or in close formation. If AOA is greater than
15 degrees, the autopilot will disconnect and theFLCS FAULT caution light and the FLCS A/P FAIL
PFL message will activate.
Additionally, the FLCC monitors autopilot operation
for a failure to maintain the selected mode and for
prolonged engagement outside of autopilot attitude
limits with no stick inputs. Detection of a failure
results in activation of the FLCS FAULT caution
light and the FLCS A/P DEGR PFL message.
Positioning the PITCH switch to ALT HOLD enables
the FLCS to use CADC information to generatecommands to the horizontal tails which result in
the aircraft maintaining a constant altitude. The
FLCS limits the pitch command to +0.5g to +2g.
Engagement of altitude hold at rates of climb or
dive less than 2000 fpm selects an altitude within thepitch command g limits. Engagement above rates of2000 fpm causes no unsafe maneuvers; however, theengaged altitude may not be captured. Control
accuracy of "100 feet is provided to 40,000 feetpressure altitude for normal cruise conditions. The
altitude reference may be changed by depressing thepaddle switch, changing altitude, and releasing thepaddle switch. ALT HOLD in the transonic region
may be erratic.
Positioning the PITCH switch to ATT HOLD allows
an attitude signal from the INU to be used tomaintain the selected pitch attitude. This mode does
not function if pitch angle exceeds "60 degrees;
however, the PITCH switch remains engaged.

1136

Positioning the ROLL switch to HDG SEL allows the
FLCS to use a signal from the HSI to maintain theheading set on the HSI. Adjusting the HSI headingreference marker to the aircraft heading prior to
engagement maintains the existing aircraft head
ing; otherwise, when the autopilot is engaged with
the ROLL switch in HDG SEL, the aircraft turns tocapture the heading indicated by the headingreference marker on the HSI. The roll command does
not exceed a 30degree bank angle or a 20degree/se
cond roll rate. This mode does not function if bankangle exceeds "60 degrees; however, the ROLL
switch remains engaged.
Positioning the ROLL switch to ATT HOLD allows
the FLCS to use an attitude signal from the INU tomaintain the aircraft at selected roll attitude. Thismode does not function if bank angle exceeds "60
degrees.
Positioning the ROLL switch to STRG SEL allows
the autopilot to steer the aircraft to the selected
steerpoint using roll commands. The roll command
does not exceed a 30degree bank angle or a 20
degree/second roll rate. This mode does not function
if bank angle exceeds "60 degrees; however, theROLL switch remains engaged.
ADV MODE Switch
ADV MODE switch is a twofunction switch. The
ACTIVE light, which is located on the top half of the
ADV MODE switch, illuminates whenever ATF is
engaged. Depression of the ADV MODE switch when
the NVP is in an operate mode engages ATF. The
STBY light, which is located on the bottom half of the
ADV MODE switch, illuminates when PITCH
autopilot is engaged and the NVP is in an operatemode. This configuration engages the TF/PITCH
autopilot blending mode; depression of the ADV
MODE switch at this point deselects the blending
mode and illuminates the ACTIVE light (ATF). The
STBY light also illuminates from the ATF mode if thepaddle switch is depressed, indicating that ATF is
ready to assume control (when paddle switch isreleased).
STICK STEERING
Stick steering is operable only with the pitch and roll
attitude hold modes. Stick steering operation isaccomplished by applying force to the stick. With
ATT HOLD selected, a force applied in theappropriate axis large enough to activate stick
steering causes the autopilot to drop the selectedreference and the system accepts manual inputs
from the stick.

T.O. GR1F16CJ1AOA SYSTEM
The system consists of two AOA transmitters locatedon each side of the nose radome, AOA ports on thefuselagemounted air data probe, a PSA, an AOA
correction device in the CADC, an AOA indexer, anda vertical scale AOA indicator. In flight, the airflow
direction is sensed by the conical AOA probes and the
AOA ports of the fuselagemounted air data probe.
The AOA signals from all three sources are sent to the
input selector/monitor of each of the four FLCC
branches.
The AOA indicator, AOA indexer, and HUD AOA
bracket display the correct AOA until NLG WOW.
Therefore, accurate AOA indication is available
throughout twopoint aerodynamic braking. WhenNLG WOW occurs, all AOA indications are based onzero degrees.

The first AOA input signal or sensing failure isindicated by the FLCS FAULT caution light and the
PFL message FLCS AOA FAIL on the PFLD and theFLCS MFD page. A second failure is indicated by the
FLCS warning light and the FLCS AOA WARN PFL
message on the PFLD. When a first AOA failure is
identified, the failed input is set to 11 degrees AOA.
If a second AOA failure subsequently occurs, the11degree signal existing in the first failed branch
prevents hardover AOA inputs and may provide AOA
information for landing.
During DBU operation, AOA failures are not reported
and a first AOA failure is not set to 11 degrees.
AOA Indicator
Refer to figure 152. The AOA indicator, located on the
instrument panel, displays actual AOA in degrees.
The indicator has a vertically moving tape display
indicating an operating range of -5 to approximately+32 degrees. The tape is color coded from 917 degrees
to coincide with the colorcoded symbols on the AOA
indexer.

AOA Displays
INDICATORINDEXERHUD DISPLAY

ATTITUDE

15SLOW
HIGH AOA

13

ON SPEEDOPTIMUM AOA

11

FAST
LOW AOA
GR1F-16CJ-1-0053X37

Figure 152.

1137

T.O. GR1F16CJ1AOA Indexer
Refer to figure 152. The AOA indexer, located on the
top left side of the glareshield, consists of threecolorcoded symbols arranged vertically. The indexer
provides a visual headup indication of aircraft AOA
by illuminating the symbols individually or in
combinations as shown. The indexer lights display
AOA correction (based on approximately 13 degrees
AOA). This correction may be used during landing
approaches as visual direction toward optimum
landing AOA. The AOA indexer operates continu
ously with the LG handle up or down.
A dimming lever, located on the left side of theindexer, controls the intensity of the lighted symbols.
The indexer lights are tested by activation of the MAL& IND LTS switch on the TEST switch panel. The test
should be performed with the dimming lever in thebright position.

Air Data Probes
Two air data probes provide data inputs to the air
data system. One air data probe (pitot probe) ismounted on the nose and provides a dual source ofstatic and total pressure. The other air data probe is
mounted on the forward right side of the fuselage andprovides a source of AOA, sideslip, static pressure,and total pressure.
AOA Transmitters
The AOA transmitters are mounted on each side of
the radome and each provides sensor data to the
FLCS proportional to local AOA. The probe of thetransmitter protrudes through the radome to alignwith the airstream.

HUD AOA Display
Refer to figure 152. The HUD AOA bracket andflightpath marker provide a visual headup indication of aircraft AOA. The flightpath marker alignedwith the top of the bracket indicates 11 degrees AOA.
The flightpath marker centered on the bracket
indicates 13 degrees AOA. The flightpath marker
aligned with the bottom of bracket indicates 15
degrees AOA. The HUD AOA display is only available
with the NLG lowered.

AIR DATA SYSTEM
Refer to figure 153. The air data system uses probes
and sensors to obtain static and total air pressures,
AOA, sideslip, and air temperature inputs. These air
data parameters are processed and supplied tovarious systems.
Proper AOA transmitter and fuselage air data probe
operation is essential for safe flight operation.
Interference from foreign objects (especially ice,internal or external) or improperly installed AOA
transmitters can result in erroneous AOA data at
weight off wheels. Reporting of false high AOA
concurrently from two sources can cause the FLCC tocommand full nose down pitch which is impossible forthe pilot to stop. Ground use of probe covers protects
the system from foreign objects and moisture
intrusion. Ice on/in the probes is eliminated by usingprobe heat prior to takeoff.

1138

Total Temperature Probe
The total temperature probe provides the CADC withan analog signal which is required for true airspeed
and air density computation. The probe is located on
the underside of the right forebody strake.
Static Pressure Ports
Two flushmounted static pressure ports used formeasuring sideslip are located on the fuselage left
and right sides aft of the forward equipment bay
doors. These two ports provide inputs to a differential
pressure sensor for angleofsideslip measurement.
The measurement is also used to compensate thethird AOA source error and to verify the AOS derived
from the INS.
Probe Heat Monitor
The probe heat monitor monitors current flow to thepitot, fuselage air data, and AOA probes (total
temperature probe current is not monitored). If thecurrent flow decreases below a certain value, themonitor illuminates the PROBE HEAT caution
light. The monitor operates anytime the aircraft isairborne, regardless of the PROBE HEAT switch
position.

T.O. GR1F16CJ1

Air Data System Schematic (Typical)
AOA TRANSMITTER
LEFT/RIGHT SIDE

FLUSH STATIC PORTRIGHT SIDEPSRSIDESLIP DIFFERENTIALPRESSURE SENSORVERTICAL VELOCITY IND

FLIGHT CONTROLCOMPUTERHEMISPHERICAL(Portion of fuselagemounted air data probe)
PS3
P
P 2
FUSELAGEMOUNTED
AIR DATA
PROBE

P1

P 2

PT3

PNEUMATICSENSORASSEMBLY(PSA)

ALTIMETER

P 1

FLUSH STATIC PORTLEFT SIDEPSLAIRSPEED/MACH IND
NOSE-MOUNTED AIRDATA (PITOT) PROBE

P T1
P T2
PS1
PS2

CAUTION LIGHT PANELAOA INDEXER
AOA

CADC

CENTRAL AIR DATACOMPUTER (CADC)LEGEND:
ELECTRICALP T = TOTAL PRESSURE
PS = STATIC PRESSURE
P
= AOA
P
= PROBE STATIC PRESSURE
PSL = STATIC PRESSURE, LEFT
PSR = STATIC PRESSURE, RIGHTTOTALTEMPERATURE
PROBE

LE FLAP COMMANDSERVOUSING SYSTEMS:
ALTIMETER
FIRE CONTROL SYSTEM(HUD AND PX II FCC, PX III MMC)AOA INDICATORIFF TRANSPONDERPROPULSION
NAVIGATIONECSFLIGHT CONTROLSAUTOPILOT
GR1F-16CJ-1-0054A37
CSFDR

Figure153.

Change 11139

T.O. GR1F16CJ1PROBE HEAT Switch
The PROBE HEAT switch is located on the TEST
switch panel. The pitot, fuselage air data, AOA, andthe total temperature probe heaters are on anytime
the aircraft is airborne, regardless of the PROBEHEAT switch position.
Functions are:
D PROBEHEAT-On the ground, this positionenergizes the pitot, fuselage air data, AOA, and thetotal temperature probe heaters and the probe heatmonitor.

D OFF - On the ground, circuits deenergized.
D TEST-On the ground and in flight, this position

performs a functional test of the probe heatmonitoring system. The PROBE HEAT caution
light flashes 35 times per second. If the caution
light does not illuminate or if it illuminates but does
not flash, the probe heat monitoring system is
inoperative. The test feature does not verify proper
operation of probe heaters.

PNEUMATIC SENSOR ASSEMBLY (PSA)
The PSA converts pneumatic inputs from the nose
air data probe and the fuselage air data probe into
electrical signals. The PSA supplies static andimpact pressure signals and single AOA signals to
the input selector/monitor of each branch of the
FLCC. A ratio of impact to static pressure is
generated within the FLCC and used with AOA andstatic pressure to schedule the LEF's and for gain
scheduling. The input selector/monitors also arecapable of detecting single and dual malfunctions of
the air data sensor signals. Single static or impact
pressure failures illuminate the FLCS FAULT
caution light and the FLCS ADC FAIL PFL on the
PFLD and MFD. A dual malfunction of static or
impact pressure signals results in the following:
D Illumination of the FLCS warning light, continued

illumination of the FLCS FAULT caution light, andthe FLCS ADC FAIL (for first failure) and STBYGAINS PFL messages.
of FLCS standby gains. Refer to
STANDBY GAINS, this section.

CENTRAL AIR DATA COMPUTER (CADC)
Refer to figure 154 for a signal flow diagram
showing the systems interacting with the CADC.
The CADC receives total and static pressures, AOA,
and total temperature inputs, converts the inputs
into digital data, and then transmits the data to theusing systems. The CADC has continuous BIT andinitiated BIT features; initiated BIT is run during
the FLCS BIT.

CADC Caution Light
The CADC caution light, located on the caution light
panel, illuminates whenever a malfunction in the
CADC is detected. If there is a mach signal failurefrom the CADC, the ENGINE FAULT caution lightalso illuminates and the ENG MACH FAIL PFL isdisplayed on the PFLD. A CADC malfunction may
result in an FLCS AOS FAIL PFL and deactivationof the AOS feedback function.

WARNING, CAUTION, AND
INDICATOR LIGHTS
Refer to figure 155. Warning, caution, and indicatorlights are used throughout the cockpit to callattention to a condition or to allow an item to be easily
read. Red warning lights and the amber MASTER
CAUTION light are all located on the edge of theglareshield. All of the lights, except the MASTER
CAUTION light, are described under their respective
systems.

D Activation

Pitot probe tip icing results in erroneously low
airspeed indications, illumination of the FLCS
FAULT caution light with an FLCS ADC FAIL PFL
message, and flight control gains scheduled for low
airspeed conditions.

1140

The warning and caution lights (except MASTER
CAUTION) are not presstotest or presstoreset
lights. Pressing these lights releases them from theirmodules and deactivates them. To reengage areleased light, pull it out slightly and then press toreengage the module.

T.O. GR1F16CJ1CADC and Interfacing Systems
STATIC PRESSURE
NOSE-MOUNTED
AIR DATA (PITOT)PROBE

TOTAL PRESSURE

TO/LDGCONFIGWARNINGDISCRETE BASED ON
CALIBRATED AIRSPEED,PRESSURE ALTITUDE, AND
PRESSURE ALTITUDE RATECENTRAL
AIR DATACOMPUTER(CADC)TOTAL TEMPERATURETOTALTEMPERATUREPROBEPRESSURE ALTITUDE (29.92)
BAROMETRIC SETTING

ALTIMETERAIMS CODEPRESSURE ALT VALIDIFF TRANSPONDER

CADC STATUS

AVIONICSYSTEMS,
PW229 EDU,AND
GE129 EMSC

AIR DENSITY RATIOPRESSURE ALTITUDE

TRUE TEMPERATURE

MACH NUMBER

PRESSURE ALTITUDE (29.92)

ENVIRONMENTALCONTROL SYSTEM

TRUE AIRSPEEDCALIBRATED AIRSPEED

PRESSURE ALTITUDE (29.92)

AOA

PRESSURE ALTITUDE (BARO)

MACH NUMBER

CALIBRATED AIRSPEED

PRESSURE ALTITUDE (BARO)TRUE TEMPERATURE

COMMAND WORD

AOA
MACH NUMBERCRASHSURVIVABLE
FLIGHT DATARECORDER(CSFDR)

PRESSURE ALTITUDE (BARO)CADC RESETFLIGHT
CONTROLCOMPUTER

115 VAC, 1

AOA L, R, 400 Hz

ELECTRICAL POWER

CADC TEST ENGAGECADC TEST PASSDISCRETESAOA INDEXER

CADC RESET
MACH NUMBER
FLIGHT CONTROL
PANEL
AOA

V V I

CADC FAIL STATUS

VERTICAL VELOCITYPROPULSIONCONTROL SYSTEMAOA INDICATORCAUTION LIGHT
PANEL

GR1F-16CJ-1-0055A37

Figure154.

1141

T.O. GR1F16CJ1

Warning, Caution, and Indicator
Lights C DF (Typical)
AOA
INDEXERDFC
THREAT WARNINGADVANCEMODE
SWITCHLG CONTROL PANELTHREAT WARNING

ENG
& JET
START
CONTROLPANEL

EPU CONTROL
PANEL

FLCP

ELEC CONTROLPANELTEST SWITCH PANEL

Figure155.(Sheet 1)

1142

ECM CONTROL PANEL

GR1F-16CJ-1-0056-1A37

T.O. GR1F16CJ1PX IIWARNING LIGHTSARE RED WITH BLACK
LETTERINGAR STATUS/NWS
INDICATORPX IIIDATA ENTRY DISPLAY

MARKER
BEACONPILOT FAULT LIST DISPLAYNWS CONTROLBUTTON/
INDICATORCAUTION LIGHT PANELPX II OXY LOWPX III OBOGS

DF

CMDS

GR1F-16CJ-1-0056-2A37

Figure155.(Sheet 2)

1143

T.O. GR1F16CJ1

Warning, Caution, and Indicator
Lights DR (Typical)
AOA
INDEXERPX II WARNING LIGHTSARE RED WITH BLACK
LETTERINGAR STATUS/NWS
INDICATORSTICKOVERRIDE
LIGHT

DATA ENTRY DISPLAYPILOT FAULT LIST DISPLAY
MARKER
BEACONCAUTION LIGHT PANELNWS CONTROLBUTTON/
INDICATORPX II OXY LOW, PX III OBOGSLG CONTROL PANEL

GR1F-16CJ-1-0057A37

Figure155. (Sheet 3)

1144

T.O. GR1F16CJ1VOICE MESSAGE SYSTEM (VMS)

D LOCKLOCK - Advises that radar has locked on to

The VMS provides a warning message, a caution
message, or discrete messages. The fixed volume
voice message does not blank other audio and,
therefore, may not be heard.

D PULLUPPULLUPPULLUPPULLUP:

The warning message (WARNINGWARNING pause
WARNINGWARNING) is automatically activated
1.5 seconds after illumination of any glareshield
mounted warning light.
The caution message (CAUTIONCAUTION) is
automatically activated 7 seconds after the illumina
tion of any light on the caution light panel except the
IFF caution light. If the MASTER CAUTION light isreset immediately after its illumination, the caution
message does not occur.
The warning/caution messages are reset for subse
quent activation by:
D Resetting the WARN RESET on the ICP for voice

warning.

D Resetting

caution.

target. Refer to T.O. GR1F16CJ3411 for a
detailed description.

S In ATF or in manual TF (flyup enabled), advisesthat an automatic flyup has been initiated.

S In manual TF (flyup disabled), advises that TF

failure, SWIMdetected failure, glimit warning,or obstacle warning has occurred.

S Advises that GAAF was activated PX III

or that
DTS ground proximity warning was activated.

D COUNTER - Active in semiautomatic mode only if

REQCTR option is turned on via the DED CMDSBINGO page. Advises that a dispense commandshould be initiated.

D CHAFFFLARE - Active if FDBK option is turned

on via the DED CMDS BINGO page. Advises that
CMDS has initiated a dispense program.

D LOW - Active if BINGO option is turned on via thethe MASTER CAUTION for voice

D Eliminating the condition that originally activatedthe lights and messages.

Discrete voice messages are provided when certain
conditions occur.
Messages are:
D ALTITUDEALTITUDE - Advises that:
S Descent is occurring after takeoff.
S Radar altitude is below the entered radar ALOW

value.

S Barometric altitude is below the entered MSL

ALOW value.

Refer to T.O. GR1F16CJ3411 for a detailed
description.
D BINGOBINGO- Advises that the bingo fuelwarning has been activated. Refer to T.O. GR1F
16CJ3411 for a detailed description.

D IFF - Not operable in flight (message is heard

during ground test).

D JAMMER - Active if REQJAM option is turned on

via the DED CMDS BINGO page. Advises that athreat should be jammed and pilot consent isrequired.

DED CMDS BINGO page. Advises that expendable
low quantity exists.

D OUT - Active if BINGO option is turned on via the

DED CMDS BINGO page. Advises that expendable
type is completely spent.

D PX III DATA-Advises that IDM received data linktarget information.

D LOCK - Not operable in flight (message is heard

during ground test).

All voice messages have priority over the low speed
warning tone and LG warning horn. Voice messages
are also prioritized.
Priority sequence is:
D PULLUP.
D ALTITUDE.
D WARNING.
D JAMMER.
D COUNTER.
D CHAFFFLARE.
D LOW.
D OUT.
D LOCK.

1145

T.O. GR1F16CJ1D CAUTION.

PILOT FAULT LIST DISPLAY (PFLD)
Refer to figure 156. The PFLD, located on the C DF
right auxiliary console, DR instrument panel,
displays engine, avionics, and FLCS PFL's. In
addition, a status line displays a system code to
identify the system associated with an active fault(s).

D BINGO.
D DATA.
D IFF.

The VMS does not function with WOW. However, it
can be tested by pressing the MAL & IND LTS test
button on the TEST switch panel. During test, eachword is heard one time. The VMS is powered bybattery bus No. 1.

VOICE MESSAGE Switch C

DF

Refer to figure 13. The VOICE MESSAGE switch,
located aft of the stick, is a twoposition switch.
Positions are marked VOICE MESSAGE and
INHIBIT. During normal operation, the switch issafetywired in VOICE MESSAGE. Placing theswitch to INHIBIT disables all voice messages.
INHIBIT should only be used to clear a voice message
which repeats abnormally. Placing the switch back to
VOICE MESSAGE enables normal operation.

Two types of PFL's are displayed: warning level andcaution level. Warning level PFL's are associated with
the FLCS (TFR system PFL's are considered to be
associated with the FLCS) and are distinguishedfrom caution level PFL's by a warning indicator whichbrackets the PFL message. When an FLCS warninglevel PFL occurs, the PFL message and FLCS codeare displayed on the PFLD, the FLCS warning lightilluminates, a flashing WARN is displayed in the
HUD, and the voice warning message is activated.
Caution level PFL's are associated with the FLCS,
engine, and avionic system. When a caution level PFL
occurs, the PFL message and system code aredisplayed on the PFLD, the appropriate system fault
caution panel light illuminates, the MASTER
CAUTION light illuminates, and the voice caution
message is activated.
If multiple PFL's occur, they are displayed in thefollowing priority order:
D FLCS warning level PFL's.
D FLCS caution level PFL's.
D Engine PFL's.

MASTER CAUTION LIGHT
The MASTER CAUTION light, located on the C DF
left upper edge, DR center of the glareshield,
illuminates shortly after an individual light on thecaution light panel illuminates (except the IFF
caution light) to indicate a malfunction or specificcondition exists.
The MASTER CAUTION light does not illuminate inconjunction with the warning lights. The MASTER
CAUTION light may be reset by depressing the faceof the light unless it is illuminated by the ELEC SYS
caution light. For FLCS FAULT, ENGINE FAULT,
and AVIONICS FAULT caution lights, the MASTER
CAUTION light may also be reset by depressing the
C DF FACK, DR FAULT ACK button. The lightshould be reset as soon as feasible so that othercaution lights can be monitored should additional
malfunctions or specific conditions occur. Unless it isreset, the MASTER CAUTION light remains
illuminated as long as the individual caution light isilluminated. DR The MASTER CAUTION light is arepeater and cannot be reset individually.
1146

D Avionic PFL's.

Each page of the PFLD displays up to three PFL's.
Additional pages are indicated by an arrow at eachend of the bottom PFL and are accessed by depressing
the C DF FACK, DR FAULT ACK button. Page
numbers are also displayed when more than three
PFL's are listed.
PFL's are acknowledged and recalled by depressing
the C DF FACK, DR FAULT ACK button.
Acknowledging a caution level PFL clears it from the
PFLD and extinguishes the associated caution panel
light and MASTER CAUTION light (if not previously
reset). Acknowledging an FLCS warning level PFL
clears it from the PFLD; however, the FLCS warning
light remains on. Subsequent depressions of the C
DF FACK, DR FAULT ACK button perform fault
recalls. At the time of a fault recall, PFL's that are nolonger being reported as a failure within theoriginating system are not displayed and are cleared
from memory. The system code (FLCS, ENG, and/or
AV) provides a real time indication of the presence ofactive fault(s). If the code is displayed, there is at least
one active PFL within the reported category. If nosystem codes are displayed, there are no active PFL's.

T.O. GR1F16CJ1

Pilot Fault List Display (Typical)
C

1

DF

2

DR

3
FLCS
FLCS
FLCS
ENG

8

7

ENG

AV

LEF
AOA
A/I

LOCK
FAIL
FAIL

6

P1

4

5

1. System Codes
2. Warning Indicator
3. Page Number
4. PFL Message(s)5. Additional Page Indicator
6. Severity Indication
7. LRU or Function
8. Subsystem
GR1F-16CJ-1-0058X37

Figure 156.

Eliminating a FLCS failure condition by corrective
action (e.g., FLCS reset) prior to fault acknowledg
ment removes all FLCS failure indications. For
engine and avionic failure conditions and for FLCS
failure conditions after fault acknowledgment,elimination of the failure condition will automaticallyremove/clear the system codes but not the PFL
messages. Fault acknowledgment (if not previously
accomplished) and fault recall are required to remove
the PFL message(s).
The PFLD blanks if the upfront controls fail. Refer toT.O. GR1F16CJ3411 for additional PFLD infor
mation. The PFLD is powered by the nacelle essential
ac bus.
CAUTION LIGHT PANEL
The caution light panel is located on the right
auxiliary console.

The ELEC SYS caution light must be reset at the
ELEC control panel with the ELEC CAUTIONRESET button. The light may appear nonresettablein situations where the ELEC SYS caution light is
rapidly flashing or cycling on and off.
The following caution lights may be reset with the C
DF FACK DR FAULT ACK button:
D FLCS FAULT.
D ENGINE FAULT.
D AVIONICS FAULT.

MAL & IND LTS Test Button
The MAL & IND LTS test button, located on the TEST
switch panel, operates relays which test theillumination of all warning, caution, and indicatorlights, the LG warning horn, and voice messages.

1147

T.O. GR1F16CJ1

LIGHTING SYSTEM

Position/Formation Lights
Refer to figure 160.

EXTERIOR LIGHTING
Refer to figure 157, 158, and 159. All of the exterior
lights except the landing/taxi lights are controlledfrom the EXT LIGHTING control panel.
Anticollision Strobe Light
The anticollision strobe light is masked to minimize
projections in the cockpit.
PX III The anticollision light has a strobe (visible tothe unaided eye) as well as an infrared emitter(covert, visible to night vision devices but invisible tothe unaided eye). The strobe is disabled and the IR
emitter enabled with the MASTER knob in COVERTALL or COVERT AC. The anticollision light can
flash seven selectable patterns in each of four power
supply switch settings (figure 159).

PX III The position/formation lights have NVIS
friendly visible color lights as well as covert IR
emitters. The visible lights are disabled and thecovert lights enabled with the MASTER knob in
COVERT ALL or COVERT FORM. The lower
formation light and the lower wingtip lights do nothave IR emitters and are disabled with the MASTER
knob in COVERT ALL or COVERT FORM.

Air Refueling Lights
The AR floodlight shares the housing of the top
fuselage formation light. The light is directed aft toflood the receptacle, fuselage, wing, and empennage.
The AR floodlight is not NVIS friendly. The AR
slipway contains embedded lights on each side of theslipway. These lights are enabled when the AIR
REFUEL switch is in OPEN. The intensity of theselights is controlled by the AERIAL REFUELINGknob. These lights are not NVIS friendly.

EXT LIGHTING Control Panel C DF
(Typical) PX II
3 4

1 2

1. ANTICOLLISION Switch
2. FLASH STEADY Switch
3. WING/TAIL Switch
4. FUSELAGE Switch
5. AERIAL REFUELING Knob
6. MASTER Switch
7. FORM Knob

765
GR1F-16CJ-1-0059A37

Figure 157.(Sheet 1)

1148

T.O. GR1F16CJ1

EXT LIGHTING Control Panel C DF
(Typical) PX IICONTROL
1.

2.

3.

4.

ANTICOLLISION Switch

FLASH STEADY Switch

WING/TAIL Switch

FUSELAGE Switch

POSITIONFUNCTIONANTICOLLISION

Turns on the white anticollision strobe (flash)light on top of vertical tailOFF

Turns off the anticollision strobe light

FLASH

Causes the lights controlled by WING/TAIL
switch to flash when turned on

STEADY

Causes the lights controlled by WING/TAIL
switch to light steady when turned on

BRT

Turns on the red and green wingtip and inlet
lights and white light at trailing edge of verti
cal tail brightOFF

Turns off the white light at trailing edge of ver
tical tail and the inlet lights. Allows the red
and green wingtip lights to be controlled by the
FORM knob

DIM

Turns on the red and green wingtip lights dim.
Turns on the red and green inlet lights and thewhite light at the trailing edge of the vertical
tail dim

BRT

Turns on the white floodlights at the base of
the vertical tail brightOFF

Turns off the white floodlights

DIM

Turns on the white floodlights dim

5.

AERIAL REFUELING
Knob

Variable from off to
BRT

Varies the AR slipway lights from off to brightif AIR REFUEL switch is in OPEN

6.

MASTER Switch

NORM

Enables all exterior lights except landing andtaxi lights

OFF

Disables all exterior lights except landing andtaxi lights

Variable from off to
BRT

Varies the white formation lights on top andbottom of fuselage and, when the WING/TAIL
switch is OFF, varies the red and green wing
tip lights from off to bright

7.

FORM Knob

Figure 157.(Sheet 2)

1149

T.O. GR1F16CJ1

EXT LIGHTING Control Panel C DF
(Typical) PX III
1 2

3 4

1. ANTI-COLLISION Knob
2. FLASH STEADY Switch
3. WING/TAIL Switch
4. FUSELAGE Switch
5. AERIAL REFUELING Knob
6. MASTER Knob
7. FORM Knob

765
GR1F-16CJ-1-1059X37

CONTROLPOSITION

6. MASTER Knob
COVERT ALL

COVERT AC

COVERT FORM

NORM

1. ANTI
1, 2, 3, 4,
COLLISION A, B, & C
K b
Knob
OFF

Flashes set pattern Flashes set pattern Flashes set pattern Flashes set pattern(covert strobe)(covert strobe)(visible strobe)(visible strobe)

2. FLASH
STEADY
SwitchFLASH

All lights flash co All lights except All lights flash co All lights except
vert
the fuselage lights vert
the fuselage lights
flash
flash

STEADYAll lights are covert All lights steadysteadyAll lights are covert All lights steadysteadyBRT

Upper wingtip, in Wing tip, intake,
take, upper forma and tail lights ontion, and tail lights visible brighton covert bright(lower wingtip andlowerformationlights off)

Upper wingtip, in Wing tip, intake,
take, upper forma and tail lights ontion, and tail lights visible brighton covert bright(lower wingtip andlowerformationlights off)

3. WING/
TAIL
SwitchOff

Off

Figure 158.(Sheet 1)1150

Off

Off

T.O. GR1F16CJ1

EXT LIGHTING Control Panel C DF
(Typical) PX IIICONTROL
3. WING/
TAIL
Switch(Cont)

4. FUSELAGE
S it hSwitchPOSITION

6. MASTER Knob
COVERT ALL

COVERT AC

COVERT FORM

NORM

OFF

Wing tip, intake
and tail lights off.
Upper formation
light covert brightIntake and taillights off. Wingtip
lights controlled
by FORM knob

Wing tip, intake
and tail lights off.
Upper formation
light covert brightIntake and taillights off. Wingtip
lights controlled
by FORM knob

DIM

Upper wingtip, in Wingtip, intake,
take, tail lights on and tail lights oncovert dim. Upper visible dim
formation light covert bright. (lower
wingtip and formation lights off)

BRT

Fuselage lights co Fuselage
lights Fuselage lights co Fuselage
lightsvert brightvisible brightvert brightvisible brightOFF

Fuselage lights off Fuselage lights off Fuselage lights off Fuselage lights off

DIM

Fuselage lights co Fuselage
lights Fuselage lights co Fuselage
lightsvert dimvisible dimvert dimvisible dim

Upper wingtip, in Wingtip, intake,
take, tail lights on and tail lights oncovert dim. Upper visible dim
formation light covert bright. (lower
wingtip and formation lights off)

5. AERIALREFUELING Knob

Variable from off to Variable from off to Variable from off to Variable from off tobrightbrightbrightbright7. FORM
Knob

No effect

Controls forma No effecttion lights (andwingtip
lightswhen WING/TAIL
switch is in OFF)variable from off
to bright

Controls formation lights (andwingtip
lightswhen WING/TAIL
switch is in OFF)variable from off
to bright

Figure 158.(Sheet 2)

1151

T.O. GR1F16CJ1

Anticollision Light Flash Patterns PX IIINOTE:Power supply settings are preset by maintenance personnel.

ANTI-COLLISION
Knob
Setting
Description

POWER SUPPLY SETTING 1OFF1

1 Flash2

2 Flash3

3 Flash4

4 FlashA

1-2 FlashB

1-3 FlashC

6 FlashANTI-COLLISION
Knob
Setting
Description

POWER SUPPLY SETTING 2OFF1

1 Flash2

2 Flash3

3 Flash4

4 FlashA

2-1 FlashB

3-1 FlashC

6 FlashGR1F-16CJ-2066X37

Figure 159.(Sheet 1)

1152

T.O. GR1F16CJ1

Anticollision Light Flash Patterns PX IIINOTE:Power supply settings are preset by maintenance personnel.

ANTI-COLLISION
Knob
Setting
Description

POWER SUPPLY SETTING 3OFF1

1 Flash2

2 Flash3

3 Flash4

4 FlashA

2-2 FlashB

2-4 FlashC

6 FlashANTI-COLLISION
Knob
Setting
Description

POWER SUPPLY SETTING 4OFF1

1 Flash2

2 Flash3

3 Flash4

4 FlashA

3-2 FlashB

4-2 FlashC

6 FlashGR1F-16CJ-1-2067X37

Figure 159.(Sheet 2)

1153

T.O. GR1F16CJ1Exterior Lighting (Typical)
POSITION/FORMATION
LIGHT (GREEN)(TOP AND BOTTOM)( PX III TOP WITH IR)(AFT SIDE)
AR (FLOOD)(WHITE)ANTICOLLISION STROBE(WHITE) ( PX III WITH IR)
VERTICAL TAIL-MOUNTEDFLOODLIGHT (WHITE)AR SLIPWAY LIGHT(INTERIOR LIGHT)

POSITION
LIGHT (WHITE)( PX III WITH IR)(BOTH SIDES)FORMATION
LIGHT (FUSELAGE LABEL ON
CONTROL PANEL)(FLOOD) (WHITE)( PX III WITH IR)(FORWARD SIDE)FORMATION LIGHT(WHITE)( PX III WITH IR)

POSITION LIGHT(R-GREEN)(L-RED)( PX III WITH IR)TAXI LIGHT
LANDING LIGHT

FORMATION LIGHT(WHITE)POSITION/FORMATION
LIGHT (RED)(TOP AND BOTTOM)( PX III TOP WITH IR)
GR1F-16CJ-1-0060A37

Figure 160.

Vertical TailMounted Floodlight
A white light is mounted on the upper leading edge ofthe vertical tail and is directed forward to flood the AR
receptacle and upper fuselage. The light illuminates
by the OPEN position of the AIR REFUEL switch.
The floodlight is not NVIS friendly.
Landing and Taxi Lights
A white landing and taxi light assembly is located on
the NLG door. The landing light is angled toilluminate the landing area. The lights are turned onby the threeposition LANDING TAXI lights switch
located on the LG control panel. The switch has
positions of LANDING, OFF, and TAXI. The lightgoes off during LG retraction if the switch is left ineither the LANDING or TAXI position. The landing
and taxi lights are not NVIS friendly.
INTERIOR LIGHTING
Refer to figure 161. The interior LIGHTING control
panel, located on the right console, contains the power
1154

and intensity controls for the primary (console andinstrument) and secondary (flood) lighting systems
for the cockpit, PX III and NVIS (night vision imaging
system). The HIGH INT position of the FLOODCONSOLES knob provides thunderstorm lighting.
PX III In addition, NVIS operation is provided whichturns off most primary and secondary lights and sets
the warning, caution, and advisory lights to dim.
The interior lighting system, with the
exception of the utility light, functions normally
except all lighting is NVIS green. Internal instru
ment lighting has been disabled and replaced withpost and bezel NVIS friendly lighting. Secondary
lighting has been equipped with NVIS green filters toeliminate IR light. Caution and warning lights havefilters which limit IR emissions to be compatible withnight vision goggles. All interior lighting (except for
one position on the utility light) is NVIS friendly andinvisible through the night vision goggles. Cockpit
instruments must be viewed beneath the goggles.
PX IIIT.O. GR1F16CJ1PX III When NVIS operation is deselected, all lightcontrol panel functions are returned to normal
operation. All console, instrument, and flood lights
are illuminated to current settings. All warning,
caution, and advisory lights return to bright only if
the PRIMARY INST knob is in HIGH INT or OFF.

PX III When the NVIS switch is in ARM, all lights
controlled by the MAL & IND LTS switch areautomatically set and locked to the dim setting andthe MAL & IND LTS switch is inoperative.

PRIMARY CONSOLES Knob

The FLOOD INST PNL knob has a cw arrow pointing
toward BRT. Rotating the knob cw varies thefloodlights intensity from off to bright. PX III During
NVIS operation, the floodlights are off and rotating
the knob has no effect on lighting intensity.

The PRIMARY CONSOLES knob has a cw arrow
pointing toward BRT. Rotating the knob cw varies theintensity of the primary console and left auxiliary
console lighting from dim to bright. PX III During
NVIS operation, this lighting is off and rotating theknob has no effect on lighting intensity.
PRIMARY INST PNL Knob
The PRIMARY INST PNL knob has a cw arrow
pointing toward BRT. Rotating the knob cw varies theintensity of the forward instrument panel and right
auxiliary console lighting from off to bright. PX III
During NVIS operation, rotating the knob cw varies
the intensity of the C DF ICP, DR IKP.
DATA ENTRY DISPLAY KnobThe DATA ENTRY DISPLAY knob has a cw arrow
pointing toward BRT. The knob controls the lightingof the DED and PFLD from dim to bright.
MAL & IND LTS Switch
The MAL & IND LTS switch has positions of BRT and
DIM and a springloaded unmarked center position.
If the PRIMARY INST PNL knob is on, momentary
activation of the switch to DIM places the lighting
system to the dim condition. The system automatical
ly returns to the BRT condition if the FLOODCONSOLES knob is turned past the detent to HIGH
INT, if the PRIMARY INST PNL knob is turned off,
or if emergency dc power is lost. The BRT condition
can be manually selected anytime. The switch
controls the light intensity of all the warning, caution,and indicator lights except the following individuallydimmed lights:
D AOA indexer.
D AR/NWS indexer.
D DED.
D ECM control panel.
D MFD's.
D PFLD.
D TWS indicators.

FLOOD INST PNL Knob

FLOOD CONSOLES Knob
The FLOOD CONSOLES knob rotates from OFF to
HIGH INT. Rotating the knob cw varies the console
floodlights intensity from off to bright. If rotated to
HIGH INT, the MAL & IND LTS and instrument
panel floodlights automatically go to bright and thealphanumeric displays controlled by the DATAENTRY DISPLAY knob go to the highest intensity
level. CCW rotation past a certain point restores thealphanumeric displays to the intensity level set by
the DATA ENTRY DISPLAY knob, but the MAL &IND LTS switch must be manually reset to DIM, if
dim is desired. PX III During NVIS operation, theconsole floodlights are off and rotating the knob has
no effect on lighting intensity.
NVIS Switch PX III
The NVIS switch has three positions: NORM LTG,
ARM, and NVIS. The ARM position is a lever locked
position. The NVIS position is spring loaded and theswitch returns to the ARM position when released.
Functions are:
D NORM LTG:
S All cockpit lighting functions normally.
S Blackout function disabled.
D ARM:S Blackout function enabled (not activated).
S Caution/warning/advisory lights function in thedim mode.

S High intensity flood lighting disabled.
S Reduces the MFD brightness level when above a

preset level.

S Deenergizes

backlighting.

the MFD nonNVIS compatible

D NVIS:
S Activates/deactivates blackout function.

1155

T.O. GR1F16CJ1Interior LIGHTING Control Panel (Typical)
1

3

2

DR

6

PX II

5

C

DF

4

1. PRIMARY CONSOLES Knob
2. PRIMARY INST PNL Knob
3. DATA ENTRY DISPLAY Knob
4. MAL & IND LTS Intensity Switch
5. FLOOD INST PNL Knob
6. FLOOD CONSOLES Knob
7. NVIS Switch
8. Hands-on Blackout (HOBO) Paddle Switch

1

3

2

4

7
8

6

5

PX IIIGR1F-16CJ-1-0061A37

Figure 161.

1156

T.O. GR1F16CJ1When the blackout function is activated, most of thenonNVIS compatible light sources in the cockpit are
turned off. These include primary flight instruments,
engine instruments, and hydraulic gauges. For acomplete list of the nonNVIS compatible lights which
are not turned off by the blackout function, refer to
GR1F16CJ3411.
HandsOn Blackout (HOBO) Switch PX III
The HOBO paddle switch, located on the throttle,
activates/deactivates the blackout function when the
NVIS switch is in ARM.
Utility Light
Refer to figure 162. The utility light, located on theright console, includes PX II three, PX III four
controls:a pushbutton switch to allow momentary
operation at the highest intensity level, an OFF DIM
BRT rotary knob to allow continuous operation at any
desired intensity level, PX III an NVIS WHITE
selector for NVIS (green) or normal (white) operation,
and a lens housing which, when rotated, adjusts thebeam from flood to spot. To release the light from its
stowed position, lower the knurled collar at the baseof the light and it will pop free. The light can be lockedback into position by placing the body of the lightparallel to the sidewall fairing and pushing down
firmly on the light assembly. With the canopy closed,
the utility light may be attached to an adjustablesliding holder located on the right body positioning
handle (towel rack). The light is powered by battery
bus No. 1.
Cockpit Spotlights
Refer to figure 162. The cockpit spotlights are locatedunder the upper left and right glareshields. In thestowed position (horizontal, facing forward), thespotlight is off. The spotlight is turned on by pullingthe spotlight barrel downward. Illumination intensity
is controlled by turning the knurled barrel (dimmer).
To turn the spotlight off, return it to the stowed
position. Overrotation of the knurled barrel may
cause breakage of the bulb or rheostat. The lights are
powered by battery bus No. 1.

ESCAPE SYSTEMCANOPYThe canopy is a twopiece, plastic, bubbletype,transparent enclosure. The forward part is a single
piece windshieldcanopy transparency, which ishinged at the aft end and is unlatched, opened, orclosed/latched by an electrically operated actuatorwith a manual backup. C A smaller fixedtransparency fairs to the fuselage aft of the seat. The
canopy may be jettisoned by internal controls forinflight or ground escape and by external controls
for ground rescue. An inflatable pressurization sealon the cockpit sill mates with the edge of the movable
canopy. A noninflatable rubber seal on the canopy
prevents the entry of water when the cockpit is notpressurized.
The canopy provides some bird strike protection. Bird
strikes on centerline at approximately eye level may
produce enough canopy deflection to shatter the HUD
combiner glass and cause rearward propagation of adeflection wave. Deflection of the canopy in the areaof the pilot's helmet has been observed to be 1 to 2
inches during bird strike tests that were consideredsuccessful. Successful completion of canopy bird
strike testing (4 pound bird at 350 or 550 knots,
depending on canopy) requires that the canopy not
deflect more than 2 1/4 inches in the area of the pilot's
helmet. This may be a consideration for adjustingseat height, especially while flying at lower altitudes
with helmetmounted equipment. Impacts off center
may not shatter the HUD glass. High energy bird
strikes may cause canopy penetration or larger
deflection waves.
CANOPY CONTROLS AND INDICATORS
Refer to figure 163.
MANUAL CANOPY CONTROL Handcrank
C DF An internal MANUAL CANOPY CONTROL
handcrank manually performs the same function as
the canopy switch. Due to the strength required toopen the canopy with the handcrank, the methodshould be considered a last resort.

An external flushmounted CANOPY handcrank
receptacle just opposite the inside manual drive isused for ground crew manual operation of the canopy.
Canopy Handle

C DF

The canopy handle, located on the canopy sill justforward of the throttle, hinges down to cover andprotect the internal canopy switch. The handle alsofunctions to inflate/deflate the canopy pressure seal,
to turn the CANOPY warning light off/on, and tomechanically prevent the canopy actuator fromunlatching. The canopy handle should be in the up(unlock) position prior to lowering the canopy.

1157

T.O. GR1F16CJ1

Utility Light and Spotlights (Typical)
PUSHBUTTON SWITCHROTARY KNOBQUICKDISCONNECT
PLUG

LENS HOUSINGCOLLAR

RIGHT BODY
POSITIONINGHANDLE(TOWEL RACK)COCKPIT SPOTLIGHTSPX IIROTARY KNOB
NVIS WHITE
SELECTORPUSHBUTTONSWITCHFLOOD SPOTLENS HOUSINGCQUICKDISCONNECT
PLUG

DF

DR

UTILITY LIGHTCOLLAR

RIGHT BODY
POSITIONINGHANDLE(TOWEL RACK)PX IIIGR1F-16CJ-1-0062A37

Figure 162.

1158

T.O. GR1F16CJ1Canopy Controls and Indicators (Typical)
NOTE:
If the internal canopy switch is in the up position,
the canopy cannot be closed using the external
canopy switch.

1(Left side panel,
forward cockpit)
CANOPYPX IIICANOPYPX II

PX IIICANOPYCANOPY

2
3

3

5

6

PX II

6

4

7 8 9

10

4

C DF Canopy Switch (internal) (spring-loaded
to center from down position)
2. C DF Canopy Handle (shown in unlocked position)
3. CANOPY Warning Light (Red)
4. Ejection Handle (PULL TO EJECT)
5. C DF MANUAL CANOPY CONTROL Handcrank
6. CANOPY JETTISON T-Handle
7. Canopy Jettison D-Handle (each side of fuselage)
8. Canopy Lock Access Plug (external)
9. CANOPY Handcrank Receptacle (external)
10. Canopy Switch (external) (spring-loaded to center)
1.

GR1F-16CJ-1-0063A37

Figure 163.(Sheet 1)

1159

T.O. GR1F16CJ1Canopy Controls and Indicators (Typical)
CONTROL/INDICATORPOSITION/INDICATIONCanopy
Switch (inter
C DFnal) (springloaded(spring loaded to centerfrom down position)

Up

Opens canopy

CenterStops canopy motion

Down

Closes and latches canopy

Up

Unlocks canopy

Down

Locks canopy

CANOPY Warning Light
(red)

Off

Canopy lockedOn

Canopy unlocked

4.

Ejection Handle (PULL TO
EJECT)

Pull

Jettisons canopy and ejects seat5.

MANUALCANOPYC DF
CONTROL Handcrank

Rotate ccw

Opens canopy

Rotate cw

Closes and latches canopy

6.

CANOPY JETTISON
THandle

Pull (depress either button)

Jettisons canopy independent of seat ejection

7.

Canopy Jettison DHandle(each side of fuselage)

Pull (approximately 6feet) (either handle)

Jettisons canopy independent of seat ejection

8.

Canopy Lock Access Plug(external)

Remove access plug

Access to unlock internal canopy handle

CANOPY Handcrank
Receptacle (external)

Rotate cw

Opens canopy

Rotate ccw

Closes and latches canopy

Canopy Switch (external)(springloaded(springloaded to center
position)

Up
C

Opens canopy

1.

2.

3.

9.

10.

Canopy
HandleC DF

Refer to EMERGENCY ENTRANCE AND
CREW RESCUE, Section III

Aft
DCenterStops canopy motion

DownC

Closes and latches canopy

Fwd
D

Figure 163.(Sheet 2)1160

FUNCTION

T.O. GR1F16CJ1CANOPY JETTISON
Pulling the external canopy jettison Dhandle,
located on either side of the fuselage, initiates the
canopy jettison sequence independent of seatejection.
Depressing the button, located on either side of theinternal CANOPY JETTISON Thandle, and pulling
the Thandle initiates the canopy jettison sequence
independent of seat ejection.ejection, the handle provides a backup to theautomatic pilot/seat separation and recovery parachute deployment system. The handle must be pulled
approximately 6 inches. The first 2 inches of pullballistically deploys the recovery parachute assem
bly; however, the seat will still be attached by thelapbelt and the inertia reel straps, and the survival
kit under the latched seat pan lid. Continued pull
releases the lapbelt and inertia reel straps andunlatches the seat pan lid.

Pulling the ejection handle (PULL TO EJECT),
located on the front of the ejection seat, initiates the
canopy jettison sequence followed by the seat ejection
sequence.

Ejection Safety Lever

EJECTION SEATSEAT NOT ARMED Caution Light

Refer to figure 164. The ACES II ejection seat is afully automatic emergency escape system. One ofthree ejection modes is automatically selected. Mode
1 is a low airspeed, low altitude mode during which
the recovery parachute assembly is deployed almost
immediately after the ejection seat departs theaircraft. Mode 2 is an intermediate airspeed, lowaltitude mode during which a drogue chute is first
deployed to slow the ejection seat followed by the
deployment of the recovery parachute assembly.
Mode 3 is a high airspeed/high altitude mode in which
the sequence of events is the same as mode 2, exceptthat automatic pilot/seat separation and deploymentof the recovery parachute assembly are delayed until
safe airspeed and altitude are reached. Controls areprovided to adjust seat height and lock shoulder
harness.
Ejection Handle
The ejection handle (PULL TO EJECT) is sized foronehanded or twohanded operation and requires apull of 4050 pounds to activate. The handle remains
attached to the seat by a wire cable after activation.
SHOULDER HARNESS Knob
The SHOULDER HARNESS knob unlocks theinertia reel when in the aft position and locks it whenin the forward position. If high longitudinal
deceleration force or high shoulder harness strap
playout velocities are encountered, the inertia reelautomatically locks.
EMERGENCY MANUAL CHUTE Handle
The EMERGENCY MANUAL CHUTE handle islocked while the ejection seat is in the aircraft. After

The ejection safety lever mechanically safeties (in the
up/vertical position) or arms (in the down/horizontal
position) the seat ejection handle.

The SEAT NOT ARMED caution light, located on thecaution light panel, illuminates when the ejection
safety lever is in the up (vertical) position. D The
caution lights illuminate independently of eachother.
SEAT ADJ Switch
Refer to figure 13. The SEAT ADJ switch is locatedon the right cockpit sidewall outboard of the stick.
Center position is springloaded off. The up position
raises the seat, while the down position lowers theseat. The seat adjustment motor is protected by athermal relay which interrupts electrical power whenoverheated. After a 1minute cooling period, themotor should operate normally. The motor is poweredby nonessential ac bus No. 1.
Shoulder Harness Straps/Parachute Risers
The upper torso restraints consist of shoulder
harness straps which also act as parachute risers.
The inertia reel straps are attached to the parachuterisers.
Inertia Reel Straps
The inertia reel straps may be manually released
after ejection by pulling the EMERGENCY MANU
AL CHUTE handle approximately 6 inches.
Lapbelt
The lower torso restraint is the lapbelt. The lapbelt
may be released after ejection by pulling the
EMERGENCY MANUAL CHUTE handle approxi
mately 6 inches.

1161

T.O. GR1F16CJ1Ejection Seat Controls and Indicators(Typical)
ENVIRONMENTAL
SENSOR PITOTS(FLIP-UP PITOTS)
INSPECTION HOLE

SAFETY PIN
HOLE

DURING EJECTION(EXTENDED PITOTS)

ALTERNATE
CONFIGURATION(FIXED PITOTS)

SIDE VIEW

INERTIA REEL STRAPSRETAINING PIN(YELLOW)RECOVERYPARACHUTEASSEMBLYENVIRONMENTAL SENSORPITOT (STOWED)RED PIN
PROTRUSION
INDICATES EXPENDEDSEQUENCERDR

INERTIA REELSTRAPS

911

C DF FLCS DATARECORDER

EMERGENCY OXYGEN HOSEQUICK-DISCONNECTSEAT NOTARMED

SEAT/HOSE
DISCONNECT

8

LAPBELT AND SURVIVALKIT STRAPSSHOULDER HARNESS STRAPS/RECOVERY PARACHUTE RISERS

710EMERGENCY OXYGEN
BOTTLE

6

5(ROTATED 180 )SURVIVAL KIT(UNDER SEATPAN LID)

4

1. Emergency Oxygen Green Ring
2. SHOULDER HARNESS Knob
3. Ejection Safety Lever
4. RADIO BEACON Switch
5. Survival KIT DEPLOYMENT Switch
6. Ejection Handle (PULL TO EJECT)
7. EMERGENCY MANUAL CHUTE Handle
8. Survival Kit Ripcord
9. Electronic Recovery Sequencer Battery Indicator
10. SEAT NOT ARMED Caution Light
11. DR EJECTION MODE SEL Handle

2
3

EMERGENCY OXYGENPRESSURE INDICATOR(ROTATED 180 )Figure 164.(Sheet 1)1162

1

LAPBELTRETAINING PIN(BOTH SIDES)(YELLOW)

GR1F-16CJ-1-0064X37

T.O. GR1F16CJ1Ejection Seat Controls and Indicators(Typical)
CONTROL/INDICATORPOSITION/INDICATION

1.

Emergency Oxygen Green
Ring

Pull

Activates emergency oxygen

2.

SHOULDER HARNESSKnobUNLOCKED (aft)Unlocks inertia reel and allows free move
ment of shoulder harness

LOCKED (forward)

Locks inertia reel and prevents forward
movement of shoulder harness

Up

Prevents pulling of ejection handle

Down

Allows ejection handle to be pulled

MAN (green dot visible)

Selects manual mode

AUTO (red dot visible)

Selects automatic mode

Survival KIT DEPLOYMENT Switch

A (forward)

Selects automatic mode

M (aft)Selects manual mode

6.

Ejection Handle (PULL TO
EJECT)

Pull

Jettisons canopy and ejects seat

7.

EMERGENCY MANUALCHUTE Handle

Pull

After ejection-ballistically deploys therecovery parachute assembly and releases
lapbelt and inertia reel straps and unlatches
the seat pan lid

8.

Survival Kit Ripcord

Pull

Deploys survival kit

9.

Electronic Recovery Se
quencer Battery Indicator

White

Sequencer thermalvated-operationalRed

Sequencer thermal batteries activated-notoperational

SEAT NOT ARMED Caution Light (amber)

On

Ejection safety lever up

OffEjection safety lever down

EJECTION
MODE SELDR
Handle (LIFT TURN)

AFT/NORM/SOLO

Selects ejection mode

3.

4.

5.

10.

11.

Ejection Safety Lever

RADIO BEACON Switch

FUNCTION

batteries

not

acti

Refer to F16D AIRCRAFT, this section

Figure 164.(Sheet 2)

1163

T.O. GR1F16CJ1Survival Kit
The survival kit is stowed under the seat pan lid. The
KIT DEPLOYMENT switch has a manual (aft) or
automatic (forward) position which selects the modeof postejection survival kit deployment. Pulling the
kit ripcord handle deploys the kit which remains
attached by a 25foot lanyard.
RADIO BEACON Switch
The RADIO BEACON switch allows the pilot to select
AUTO or MAN. In AUTO (red dot visible), the beacon
activates after pilot/seat separation. In MAN (green
dot visible), the beacon does not activate. The beacon
may be activated when on the ground if the RADIOBEACON switch is placed to AUTO or it can beremoved and manually operated as desired. The
beacon transmits on 243.0 MHz.
Emergency Oxygen
Emergency oxygen supply is automatically activatedduring ejection or may be manually activated bypulling the emergency oxygen green ring.

If the automatic pilot/seat separation and recovery
parachute deployment system fails, pulling the
EMERGENCY MANUAL CHUTE handle approxi
mately 6 inches ballistically deploys the recovery
parachute assembly and releases the lapbelt andinertia reel straps and unlatches the seat pan lid.
The liferaft, survival kit, and radio beacon antenna
are deployed following pilot/seat separation when thesurvival KIT DEPLOYMENT switch is in AUTO. If
the parachute is equipped with SEAWARS, theparachute risers are automatically released within 2
seconds after entering saltwater.
Seat ejection also automatically performs an escapezeroize operation by purging coded electronic
information associated with the following equipment:
D DTC.
D PX II FCC.
D PX III MMC.
D PX III EGI.
D PX II GPS.
D AIFF mode 4.
D PX III IDM.
D MFDS.

EJECTION SEAT OPERATION
Seat ejection is initiated by pulling the ejection
handle (PULL TO EJECT). This action retracts theshoulder harness straps and locks the inertia reel,fires the initiators for canopy jettison, and ignites two
canopy removal rockets. As the canopy leaves theaircraft, lanyards fire two seat ejection initiators.
A rocket catapult propels the seat from the cockpit
exposing the seat environmental sensor pitots andactivating the emergency oxygen. The recovery
sequencer selects the correct ejection mode, ignites
the stabilization package (STAPAC) rocket and the
trajectory divergence rocket, and (if in mode 2 or 3)
initiates the drogue gun.

1164

D TWS.

EJECTION MODE ENVELOPES
Refer to figure 165.
EJECTION SEQUENCE TIMES
Refer to figure 166.
CANOPY JETTISON/SEAT EJECTION
Refer to figure 167.
EJECTION SEAT PERFORMANCE
Refer to figure 168.

T.O. GR1F16CJ1PRESSURE ALTITUDE

1000 FEET

Ejection Mode Envelopes

16

12

MODE 38

MODE 2
MODE 1

4

0
0

100

200

300
400
500
AIRSPEED KIAS

600

700

1F-16X-1-0011X

Figure 165.

Ejection Sequence Times
NOTE:

TIME (SECONDS)EVENT

D In mode 3, events after drogue deployment
are delayed until within mode 2 envelope.
Recovery parachute deploys 1 second after
entering mode 2 envelope.
D
D Times
in the aft/forward sequenceincrease to include a 0.33second delay forthe rear seat and a 0.73second delay forthe forward seat. In SOLO, the forward seatis delayed 0.33 second.
D Canopy jettison time varies from 0.75
second at 0 KIAS to 0.13 second at 600 KIAS.
Ejection begins when canopy jettison initi
ates seat lanyards.

Mode 1Mode 2

1.

Catapult Initiation

0.0

0.0

2.

Drogue Gun Fired

NA

0.17

3.

Drogue Chute Inflated

NA

0.38

4.

Parachute Fired

0.20

1.17

5.

Seat/Drogue Separation

NA

1.32

6.

Pilot/Seat Separation

0.45

1.42

7.

Recovery Parachute Inflated

1.80

2.80

8.

Survival Kit Deployed

5.50

6.30

Figure 166.

1165

MODE 1

T.O. GR1F16CJ1

1166

Canopy Jettison/Seat Ejection
MODE 2
7 T = 2.8
7 T = 1.8
6 T = 1.42
6 T = 0.45

5 T = 1.32

4 T = 1.17

3 T = 0.38
4 T = 0.20

2 T = 0.17

8 T = 6.3
1 T = 0

1 T = 08 T = 5.5

1F-16X-1-0012X

Figure 167.

T.O. GR1F16CJ1Ejection Seat Performance C
MINIMUM EJECTION ALTITUDE vs DIVE ANGLE AND SPEEDNOTES:
ZERO PILOT REACTION TIME
WINGS LEVELSEA LEVEL
95th PERCENTILE PILOTMINIMUM EJECTION ALTITUDE AGL

FEET

1600
1400

UPRIGHT
1200
1000
800
600
400
200

MODE 1
MODE 20

MINIMUM EJECTION ALTITUDE AGL

FEET

1600
1400

INVERTED
1200
1000
800
600
400
200

MODE 1
MODE 2

0
0

15

30

45

DIVE ANGLE

60

75

90

DEGREES
1F-16X-1-2002X

Figure 168.(Sheet 1)

1167

T.O. GR1F16CJ1Ejection Seat Performance C
MINIMUM EJECTION ALTITUDE vs SINK RATE AND SPEEDNOTES:
ZERO PILOT REACTION TIME
WINGS LEVELSEA LEVEL
95th PERCENTILE PILOT600

MINIMUM EJECTION ALTITUDE AGL

FEET

500

INVERTED
400

300

200

UPRIGHT
100

MODE 1
MODE 2
0
0

2000

4000
SINK RATE

6000

8000

10,000

FEET PER MINUTE
1F-16X-1-2003X

Figure 168.(Sheet 2)

1168

T.O. GR1F16CJ1Ejection Seat Performance C
MINIMUM EJECTION ALTITUDE vs SPEED, DIVE ANGLE,
AND BANK ANGLENOTES:
ZERO PILOT REACTION TIMESEA LEVEL
95th PERCENTILE PILOT

1600

1400

MINIMUM EJECTION ALTITUDE AGL

FEET

1200

1000

800

600

400

200

MODE 1
MODE 2
0
0

15

30

45

DIVE ANGLE

60

75

90

DEGREES
1F-16X-1-2004X

Figure 168.(Sheet 3)

1169

T.O. GR1F16CJ1Ejection Seat Performance D
MINIMUM EJECTION ALTITUDE vs DIVE ANGLE AND SPEEDNOTES:
ZERO PILOT REACTION TIME
WINGS LEVELSEA LEVELFORWARD SEAT DURING DUAL,SEQUENCED EJECTION
95th PERCENTILE PILOT

2800

MINIMUM EJECTION
ALTITUDE AGL
FEET

2400

UPRIGHT
2000
1600
1200
800
400

MODE 1
MODE 202800

MINIMUM EJECTION
ALTITUDE AGL
FEET

2400

INVERTED
2000
1600
1200
800
400

MODE 1
MODE 2

0
0

15

30

45

DIVE ANGLE

60

75

90

DEGREES
1F-16X-1-2005X

Figure 168.(Sheet 4)1170

T.O. GR1F16CJ1Ejection Seat Performance D
MINIMUM EJECTION ALTITUDE vs SINK RATE AND SPEEDNOTES:
ZERO PILOT REACTION TIME
WINGS LEVELSEA LEVELFORWARD SEAT DURING DUAL,SEQUENCED EJECTION
95th PERCENTILE PILOT

1200

FEET

1000

MINIMUM EJECTION ALTITUDE AGL

800

600

INVERTED
400

200

UPRIGHT

MODE 1
MODE 2
0
0

2000

4000
SINK RATE

6000

8000

10,000

FEET PER MINUTE
1F-16X-1-2006X

Figure 168.(Sheet 5)

1171

T.O. GR1F16CJ1Ejection Seat Performance D
MINIMUM EJECTION ALTITUDE vs SPEED, DIVE ANGLE,
AND BANK ANGLENOTES:
ZERO PILOT REACTION TIMESEA LEVELFORWARD SEAT DURING DUAL,SEQUENCED EJECTION
95th PERCENTILE PILOT

2800

2400

MINIMUM EJECTION ALTITUDE AGL

FEET

2000

1600

1200

800

400

MODE 1
MODE 2
0
0

15

30

45
DIVE ANGLE60
DEGREES

Figure 168.(Sheet 6)1172

75

90
1F-16X-1-2007X

T.O. GR1F16CJ1OXYGEN SYSTEM PX III

EMERGENCY OXYGEN SUPPLY (EOS)

The oxygen system consists of an onboard oxygen
generating system (OBOGS), a backup oxygen supply(BOS), and a regulated 50 cubic inch emergency
oxygen supply (EOS) bottle.

The emergency oxygen supply (EOS) consists of ahighpressure bottle mounted on the left side of theejection seat. The EOS is a regulated oxygen supply
that will last for 8 to 12 minutes depending on
altitude and breathing rate. The hose is routed to theright side of the seat. The system is activated:

ONBOARD OXYGEN GENERATING SYSTEM(OBOGS)The OBOGS is a molecular sieve/concentrator thatuses an adsorption filtering process to remove
nitrogen from the ECS supplied bleed air and provide
oxygen rich breathing gas. The system is capable ofproducing 95 percent oxygen with the balance beingargon and other inert gases. For normal operation theconcentrator provides oxygen to a compatible diluter
demand breathing regulator. The system pressure to
the regulator can be read on a pressure gage locatedon the face of the regulator. The normal operatingpressure range is 2540 psi.
OBOGS BACKUP OXYGEN SUPPLY (BOS)
If the OBOGS is not producing oxygen due to loss of
ECS input air to the concentrator, concentrator
failure, or loss of power to the concentrator, thesystem will revert to stored oxygen in the sieve filledplenum, called the backup oxygen supply (BOS). The
BOS will last for a period of 3 to 5 minutes dependingon altitude and breathing rate. Resupplying adepleted BOS requires up to 10 minutes. After the
BOS has been expended and system pressure falls
below 5 psi, the OXY LOW warning light, located on
the right glare shield, illuminates.
OBOGS MONITORThe OBOGS is monitored by an oxygen monitor withbuiltintest (BIT) circuitry. The oxygen monitor
samples oxygen for the correct oxygen concentration
at a particular altitude from 10,000 to 50,000 feet.
The OBOGS monitor provides a signal to turn on the
OXY LOW warning light if the partial pressure ofoxygen (PPO2) from the concentrator drops below apreset alarm range or if BIT detects an internalmonitor fault. The monitor provides low oxygen
partial pressure warning at either of two thresholds
based on the position of the diluter lever located on
the regulator. Low PPO2 can be caused by aconcentrator malfunction or a clogged input filter on
the concentrator.

D Automatically upon ejection.
D Manually by pulling the emergency oxygen green

ring located on the left aft side of the seat.

A gauge on the EOS indicates the stored oxygenpressure.
OBOGS BUILTINTESTS (BIT's)
The OBOGS monitor has four BIT programs:
powerup BIT, periodic BIT, manual BIT, andpneumatic BIT. All BIT programs are for testing themonitor's circuitry only.
OBOGS PowerUp BIT
Powerup BIT is performed automatically upon
powerup during engine start. The powerup BIT
includes a two minute warmup period during which
the OXY LOW warning light is inhibited. The OXY
LOW warning light illuminates at the end of thepowerup BIT if a monitor failure is detected.
OBOGS Periodic BIT
Periodic BIT is performed automatically during
normal operations. If a monitor fault is detected, theOXY LOW warning light will illuminate.
OBOGS Manual BIT
Manual BIT is activated by momentarily positioning
the OBOGS BIT switch, located on the TEST panel,to BIT. This can be accomplished to perform a monitor
check or to determine the cause of an OXY LOW
warning light.
Selecting BIT causes the OXY LOW warning light toilluminate for ten seconds. If no faults are detected,
the warning light will go off. If a monitor fault is
detected, the OXY LOW warning light will remain onwith a continuous indication as long as the fault
condition exists.

1173

T.O. GR1F16CJ1
If the OBOGS BIT switch is positioned to BIT while
the OXY LOW warning light is on, the warning lightwill remain on steady if the problem is due to a low
PPO2 condition or low regulator pressure. It will flash
one second on and one second off if a monitor fault was
detected. If a monitor fault exists, subsequent
positioning of the OBOGS BIT switch to BIT will
toggle the warning light from flashing to continuous.

and reduce fatigue. Air pressure from the antig valve
is used by the oxygen regulator to control the amountof pressurization supplied to the oxygen mask,
helmet, and vest. A malfunction in the OBOGS
system can result in a degradation in PBG system
performance.
OXYGEN SYSTEM CONTROLS AND INDICATORSOBOGS Pneumatic BIT
Refer to figure 169 for a description of the oxygen
system controls and indicators.

Pneumatic bit is a maintenance function.
PRESSURE BREATHING FOR G (PBG)
The PBG mode of the oxygen regulator provides
pressure breathing above 4g's to enhance g tolerance

OXYGEN SYSTEM SCHEMATIC
Refer to figure 170.

Oxygen System Controls and Indicators
2
3(Typical) PX III
1

4

8
9

7
1. C DF CABIN PRESS ALT
2. OXYGEN REGULATOR FLOW Indicator
3. OXYGEN SUPPLY Indicator
4. OBOGS BIT Switch
5. OBOGS Caution Light
6. Supply (Mode) Lever (PBG Lever Lock)
7. Diluter Lever
8. EMERGENCY Lever
9. OXY LOW Warning Light

Figure 169.(Sheet 1)1174

6
5

GR1F-16CJ-1-1073X37

T.O. GR1F16CJ1Oxygen System Controls and Indicators
PX III (Typical)
CONTROL/INDICATORPOSITION/INDICATION

1.

C DF
CABIN
PRESS ALT

050,000 feet

Indicates cockpit altitude2.

OXYGEN Regulator FLOW
Indicator

White

Indicates oxygen flow

Black

Indicates no oxygen flow

3.

OXYGEN SUPPLY Indica
torOxygen pressure (psi)Indicates gaseous oxygen pressure at regula
tor in psi

4.

OBOGS BIT Switch

BIT

Initiates OBOGS manual BIT

5.

OBOGS Caution Light (am
ber)

On

Indicates that ECS air pressure has dropped
below 10 psi

6.

SUPPLY (Mode) Lever

PBG (lever lock)

Provides oxygen supply to mask, helmet
bladder, and vest. Pressure breathing as afunction of g is available

ON

Provides oxygen supply to mask, helmet
bladder, and vest. Pressure breathing as afunction of g is not available

OFF

Turns off oxygen supply

NORM

Provides regulated mixture of cockpit air andoxygen as determined by cockpit altitude

100%

Provides maximum amount of oxygen

NORM

Provides normal operation. Positive pressure
is provided if cockpit altitude exceeds
28,00032,000 feet

EMERGENCY

Provides maximum amount of oxygen underpositive pressure. This position is used by thepilot to test for leaks

TEST MASK

Provides maximum amount of oxygen underpositive pressure. This position is used fortesting by life support maintenanceOn

Indicates that partial pressure of oxygen(PPO2) is low, the presence of a BIT detected
monitor fault, or that regulator pressure has
dropped below 5 psi

7.

8.

9.

Diluter Lever

EMERGENCY Lever

OXY LOW Warning Light(red)

FUNCTION

Figure 169.(Sheet 2)

1175

T.O. GR1F16CJ1Oxygen System Schematic (Typical) PX III
CAUTION
LIGHT

OBOGS

OXYGEN MASKHELMET
BLADDERCONNECTORWARNINGLIGHT

OXY LOWOBOGSMONITORPRESSURE SIGNALFROM ANTI-G VALVEHARNESSMOUNTEDCONNECTOR

VEST
CONNECTOREOS REGULATORPRESSURE
SWITCHSEAT TO
AIRCRAFTS

VEST FILL/DUMP
VALVEPRESSURE RELIEF VALVE(FUNCTIONAL WHEN VEST
IS NOT CONNECTED)OXYGEN REGULATOR PANELPILOT TO SEATON VALVE WITH
EMERGENCY OXYGENGREEN RING

EMERGENCY OXYGEN SUPPLY(EOS) BOTTLE
EMERGENCY OXYGENPRESSURE INDICATOR

BACK-UP OXYGENSUPPLY (BOS) PLENUM

PUMP MOTORFILTER ASSY

LEGEND:
ELECTRICAL
FILL CONNECTORQUICK DISCONNECT
CONCENTRATORCHECK VALVEECS AIRGASEOUS OXYGENECS LOW PRESSURE
SWITCHECS AIR SUPPLY

GR1F-16CJ-1-1074X37

Figure 170.

1176

T.O. GR1F16CJ1OXYGEN SYSTEM PX II

Oxygen System Controls and Indicators

The oxygen system consists of a liquid system and an
emergency gaseous system.

Refer to figure 171 for description of the oxygen
system controls and indicators.
Oxygen Duration

LIQUID OXYGEN SYSTEM
A 5liter liquid oxygen system provides breathingoxygen to a diluter demand oxygen regulator. The
regulator provides for selection of normal diluted
oxygen and 100 percent oxygen and selection of PBG.
Quickdisconnects are used to expedite egressing theaircraft on the ground. Oxygen duration varies
depending upon altitude, regulator settings, andusage.

Refer to figure 172.
EMERGENCY OXYGEN SYSTEM
The emergency oxygen system consists of a highpressure bottle with a flow controller mounted on theleft side of the ejection seat. The hose is routed to theright side of the seat. The system is activated:
D Automatically upon ejection.

Pressure Breathing for G (PBG)
The PBG mode of the oxygen regulator provides
pressure breathing above 4g's to enhance g tolerance
and reduce pilot fatigue. ECS air used to inflate thegsuit is also used by the oxygen regulator to control
the amount of pressurization supplied to the oxygenmask, helmet, and vest.

D Manually by pulling the emergency oxygen green

ring located on the left aft side of the seat.

OXYGEN SYSTEM SCHEMATIC
Refer to figure 173.

Oxygen System Controls and Indicators
2
3(Typical) PX II

7

1

6
9

5

4

1. C DF CABIN PRESS ALT
2. OXYGEN REGULATOR FLOW Indicator
3. OXYGEN SUPPLY Indicator
4. Mode Lever (PBG Lever Lock)
5. Diluter Lever
6. EMER Lever
7. C DF LIQUID OXYGEN Quantity Indicator
8. C DF OXY QTY Indicator Test Switch
9. OXY LOW Caution Light

8

GR1F-16CJ-1-0073A37

Figure 171.(Sheet 1)

1177

T.O. GR1F16CJ1Oxygen System Controls and Indicators(Typical) PX IICONTROL/INDICATORPOSITION/INDICATION

1.

C DF
CABIN
PRESS ALT

050,000 feet

Indicates cockpit pressure altitude2.

OXYGEN REGULATORFLOW Indicator

White

Indicates oxygen flow

Black

Indicates no oxygen flow

3.

OXYGEN SUPPLY Indica
torOxygen pressure (psi)Indicates gaseous oxygen pressure at regula
tor in psi

4.

Mode Lever

PBG (lever lock)

Provides oxygen supply to mask, helmet
bladder, and vest. Pressure breathing as afunction of g is available

ON

Provides oxygen supply to mask, helmet
bladder, and vest. Pressure breathing as afunction of g is not available

OFF

Turns off oxygen supply

NORM

Provides regulated mixture of cockpit air andoxygen as determined by cockpit pressure
altitude100% O2

Provides regulated 100 percent oxygen

NORM

Provides normal operation. Positive pressure
is provided if cabin pressure altitude exceeds
28,00032,000 feet

EMER

Provides 100 percent oxygen under positive
pressure. This position is used by the pilot totest for leaks

TEST MASK

Provides 100 percent oxygen under positive
pressure. This position is used for testing bylife support maintenance5.

6.

Diluter Lever

EMER Lever

Figure 171.(Sheet 2)1178

FUNCTION

T.O. GR1F16CJ1Oxygen System Controls and Indicators(Typical) PX IICONTROL/INDICATORPOSITION/INDICATIONFUNCTION

7.

C DF LIQUID OXYGEN
Quantity Indicator

05 liters

Indicates quantity of liquid oxygen

8.

C DF OXY QTY Indicator
Test Switch

OXY QTY

Drives pointer ccw toward 0. As pointerpasses 0.5 liter, OXY LOW caution lightcomes on. Light goes off after the switch isreleased and the pointer passes 0.5 liter

9.

OXY LOW Caution Light(amber)On

Indicates oxygen quantity is less than 0.5
liter or system pressure below 42 psi

Figure 171.(Sheet 3)

1179

T.O. GR1F16CJ1Oxygen Duration PX IIOXYGEN DURATION (HOURS)COCKPITPRESSURE
ALTITUDE(FEET)

DILUTERLEVER(POSITION)

35,000AND UP

30,00030 00025,00020,00015,00010,000

INDICATED QUANTITY (LITERS)
5

4

3

2

1

100%

30.94

24.75

18.56

12.37

6.19

NORM

30.94

24.75

18.56

12.37

6.19

100%

22.63

18.11

13.58

9.05

4.53

NORM

23.00

18.40

13.80

9.20

4.60

100%

17.48

13.98

10.49

6.99

3.50

NORM

21.72

17.37

13.03

8.69

4.34

100%

13.19

10.55

7.91

5.28

2.64

NORM

24.43

19.55

14.66

9.77

4.89

100%

10.62

8.49

6.37

4.25

2.12

NORM

29.86

23.89

17.92

11.94

5.97

100%

8.53

6.83

5.12

3.41

1.71

NORM

29.86

23.89

17.92

11.94

5.97

NOTES:
1. Oxygen duration increases as cockpit pressure altitude increases because there is less ambient pressure
acting upon the lungs. Therefore, a smaller quantity of oxygen at altitude expands the lungs to the same
size that they were at sea level.
2. With two people, decrease NORM/100% oxygen duration by 50 percent.
3. Use of PBG may reduce the listed available oxygen times.

Figure 172.

1180

T.O. GR1F16CJ1Oxygen System Schematic (Typical) PX IIOXYLOWCAUTIONLIGHT

OXYGEN MASKPRESSURE SIGNALFROM ANTI-G VALVEHARNESSMOUNTEDCONNECTORQUANTITY
INDICATORCDFHELMET
BLADDERCONNECTORPRESSURE
SWITCHVEST
CONNECTORSEAT TO
AIRCRAFTS

VEST FILL/DUMP
VALVEPRESSURE RELIEF VALVE(FUNCTIONAL WHEN VEST
IS NOT CONNECTED)OXYGEN REGULATOR PANELCPILOT TO SEATDFDRON VALVE WITH
EMERGENCY OXYGENGREEN RING

EMERGENCY OXYGEN BOTTLE
EMERGENCY OXYGENPRESSURE INDICATORHEATEXCHANGERRELIEF VALVEOVERBOARDVENT

SUPPLYVENT

PRESSURE
CLOSING
VALVE

FILL/BUILDUP/
VENT VALVELEGEND:
FILL CONNECTORQUICK DISCONNECT

5-LITERLIQUIDOXYGENRSVRCHECK VALVEELECTRICAL
LIQUID OXYGENGASEOUS OXYGENBUILDUP COIL
LOX CONVERTER (REMOVABLE)

GR1F-16CJ-1-0074A37

Figure 173.

1181

T.O. GR1F16CJ1COMMUNICATIONS, NAVIGATION,
AND IFF (CNI) EQUIPMENT
Cockpit controls for CNI equipment are divided
between control units on the consoles and the upfront
controls located on the instrument panel. Controls forless frequently used functions, such as power andaudio volume, and essential functions, such ascommunications backup and guard, are located on
console panels. Controls for frequently used functionsof CNI are located on the upfront controls to permit
headup control during flight.
UPFRONT CONTROLS
Refer to figure 174. The upfront controls provide a
simplified, centralized, headup means of controllingthe most frequently used functions of the communica
tions system, navigation systems, and IFF. The
upfront control set consists of the data entry display(DED), C DF the integrated control panel (ICP), andDR the integrated keyboard panel (IKP). The upfront
controls are powered by emergency dc bus No.2 andemergency ac bus No. 2. Refer to F16D AIRCRAFT,
this section, for supplementary D information.

DED/CNI READOUTS
Refer to figure 174, 175, and 176. The DED is an
integral part of the upfront controls and provides avisual display of the switch actions made via upfront
controls. The primary readouts of communication,navigation, and IFF systems are included in the page
selections available for display on the DED. Channel,
frequency, mode, and code selections of UHF, VHF,
TACAN, ILS and IFF are presented when theappropriate page is selected. Present position andsteerpoint data may be selected for display on the
DED.
Displays for UHF, TACAN, and IFF show the word
BACKUP when the PX II CNI, PX III C & I knob is in
BACKUP. The word GUARD appears on UHF and
VHF pages when guard is selected.

Upfront Controls (Typical)
C

DF

DR
DATA ENTRYDISPLAY (DED)

1

2

3

4 5

6

11

1
1. COM 1 (UHF) Override Button
2. COM 2 (VHF) Override Button
3. IFF Override Button
4. LIST Override Button
5. RCL Button
6. ENTR Button
7. M-SEL Button
8. WARN RESET Switch
9. Data Control Switch (DCS)
10. Increment/Decrement Switch
11. T-ILS Function Button

INTEGRATEDCONTROLPANEL (ICP)

10

9

8

Figure 174.

1182

11

3

4

5
6
7

INTEGRATEDKEYBOARD
PANEL (IKP)

10

7

2

9

GR1F-16CJ-1-0075X37

T.O. GR1F16CJ1

CNI Readout/DED Page Summary

Backup or Guard Pages (Typical)

Normal Display Pages (Upfront Control)
UHF

STPT

10

5

VHF

12

10 : 32 : 47

M1

34CS 4567

STBY

PX II

DISPLAYED ON INITIAL POWER-UP

T123X

UHF
BUP
VHF
GRD
M

STPT
12

5

10 : 32 : 47
BUP

CNI PAGE
UHF
10
PRE

MAIN

UHF
COM
1

* 235.00 *
TOD
NB

10
226.00ON

BACKUP
10 235.00

UHF PAGE
HQ-TNG
002
PRE

12
123

All 3
Positions
Function

TNET
12 *
TOD

*

COM
1

M-SEL
0-

ENTR

RTN

SEQ

HAVE QUICK PAGE
12PRE

VHF ON12126.00VHF

COM
2* 135.00 *ON* AM * GUARDNB

VHF PAGETCN

T/R

ILS ONTCN ON

T-ILS
1* CMD STRG
24
FRQ 106.10
X (0) CRS 10

*
CHAN
BAND

BACKUPILS ONFRQCRS

CMD STRG
106.10205

TCN/ILS PAGEIFF

IFF ONMAN

M1 25
M2 1432
M3 2365

MC (5 ) **M4 A (6 )OUT (7 ) MS (8)IFF

MAN PAGEIFF ON

AUTO

M1
M2
M3

MC
M4

ALL IFF MODES

POS
(5 )

ON

BACKUP

1

*
*
NOF04
MS (8)POS PAGEIFF

IFF ON

AUTO

M1 21
M2
M3 1234

MC (5 ) **M4 A(6 ) 01:36
MS

TIM

RTN

SEQ

2

TIM PAGE

GR1F-16CJ-1-0076A37

Figure 175.

1183

T.O. GR1F16CJ1

CNI Readout/DED Page Summary PX III

Backup or Guard Pages (Typical)

Normal Display Pages (Upfront Control)
UHF

STPT

10

5

VHF

12

10 : 32 : 47

M1

34CS 4567

STBY

DISPLAYED ON INITIAL POWER-UP

T123X

UHF
BUP
VHF
GRD
M

STPT
12

5

10 : 32 : 47
BUP

CNI PAGE
UHF
10
PRE

BOTH

UHF
COM
1

* 235.00 *
TOD
NB

10
226.00ON

BACKUP
10 235.00

UHF PAGE
HQ-TNG
002 TNET
PRE

11
003

All 3
Positions
Function

BOTH
12 *
TOD

*

COM
1

M-SEL
0-

ENTR

RTN

SEQ

HAVE QUICK PAGE
12PRE

VHF ON12126.00VHF

COM
2* 135.00 *ON* AM * GUARDNB

VHF PAGETCN

T/R

ILS ONTCN ON

T-ILS
1* CMD STRG
24
FRQ 106.10
X (0) CRS 10

*
CHAN
BAND

BACKUPILS ONFRQCRS

CMD STRG
106.10205

TCN/ILS PAGEIFF

IFF ONMAN

M1 25
M2 1432
M3 2365

M4 A (6 ) **MC (5 )AUD (7 ) MS (8)IFF

MAN PAGEIFF ON

AUTO

M1
M2
M3

M4
MC

ALL IFF MODES

POS

1

*
*
NOF04
MS (8)(5 )POS PAGEIFF

IFF ON

AUTO

M1 21
M2
M3 1234

M4 B (6 ) *
*
MC
01:36
MS

TIM

RTN

SEQ

4

TIM PAGE

GR1F-16CJ-1-1076X37

Figure 176.

1184ON

BACKUP

T.O. GR1F16CJ1
The current state of CNI operation is shown on the
DED and is automatically displayed on initial power
up of the upfront controls or may be selected bypositioning the data control switch (DCS) to RTN
momentarily.
CNI Readout/DED Page Summary
The CNI page displays data concerning the UHF(COM 1), VHF (COM 2), steerpoints, IFF, systemtime, and TACAN. The CNI page displays its data on
five lines.
The first line displays information concerning the
UHF radio mode, preset channel or manualfrequency, and current steerpoint number. The label
at the far left indicates the radio mode (either UHF
or HQ). UHF is displayed when the normal mode isselected; HQ is displayed when the AJ mode isselected. The UHF/HQ label is highlighted when the
UHF VHF IFF transmit switch is keyed to UHF.
Adjacent to the mode label is the preset UHFchannel, if one is selected. If a preset channel is notselected, this area contains the manual frequency.
The right side of the first line displays navigation
information. If a steerpoint is selected, STPT isdisplayed. If an initial point is selected, IP isdisplayed. Adjacent to the steerpoint label is thenumber of the current steerpoint. Line one displays
UHF (normal mode), the selected preset UHFchannel (10), and the selected steerpoint (STPT 5).
The left side of the second line displays the status of
the UHF radio and default value indication. During
normal operation, this area is blank. OFF, BUP, and
GRD are other indications of UHF radio status. OFF
is displayed when the power is turned off with the
COMM 1 power knob. When BACKUP or GUARD is
selected, BUP or GRD is displayed. The middle areaof line two is blank during normal operation. If the
upfront controls memory is lost, the upfront controls
revert to default values and DFLT is displayed in the
middle of line two. DFLT is displayed if UFC memory
is lost and the UFC automatically uses default
values. The right side of the second line can display
wind direction and velocity. When the SEQ positionof the DCS is activated, wind direction (referenced tomagnetic north) and velocity, in knots, is displayed.
When the SEQ position of the DCS is again
activated, wind data is blanked. If wind data isdisplayed when the CNI page is exited, wind data isdisplayed when the CNI page is reentered . Wind
data is not displayed upon UFC powerup.
The third line displays VHF (COM 2) function,
preset channel or manual frequency, and systemtime. The left side of line three always displays theradio mode label VHF. When the UHF VHF IFF
transmit switch is keyed to VHF, the VHF label ishighlighted. Adjacent to the VHF label is the active
preset channel or manual frequency number. The
right side of line three shows VHF, active preset VHF
channel (12), and system time.
The fourth line displays the status of the VHF radio.
A blank area indicates normal operation. OFF
indicates that the power is off, and GRD indicatesthat GUARD has been selected. The CNI readout/
DED page summary displays blank for normal
operation. The right side of the fourth line also
displays hack time anytime the stopwatch isactivated.
The fifth line displays IFF and TACAN information.
The area at the far left displays IFF modes selected.
M, which stands for mode, is always displayed; 1, 2,
3/A, 4, C, and S are displayed when thecorresponding mode is selected. When a mode 4 reply
is issued, the 4 is highlighted. The next area of line
five displays the fourdigit code associated with IFF
mode 3/A. The next area displays IFF status. During
normal operation, this area is blank. OFF, STBY,
BUP, and AUTO are other indications of IFF status.
OFF is displayed when the power is turned off with
the IFF MASTER knob. When STANDBY orBACKUP is selected, STBY or BUP is displayed.
AUTO is displayed when the automatic transponderswitching function of IFF is selected. The right side
displays TACAN channel number and band. TACAN
channel number and band are not displayed when
TACAN radio is off or in backup. The CNI
readout/DED page summary displays IFF modes 1,
3/A, 4, C, and S selected, 4567 (mode 3/A code),
AUTO (automatic transponder switching function),
T123 (TACAN channel number), and X (TACANband).
To change preset channels or steerpoint numbers,
the up/down triangles must be next to the item to bechanged. When the DCS is positioned up or down,
the up/down triangles move between the steerpoint
and the UHF and VHF preset channels. With theup/down triangles next to the desired item, thesteerpoint/channel can be changed by depressing theincrement/decrement switch until the desired
steerpoint or channel is selected. If a manualfrequency is selected, the triangles are displayed but
are not functional. The up/down triangles are
displayed at the steerpoint number during initial
upfront control powerup. UHF, VHF, IFF, TACAN,
and ILS detailed operation using the upfront
controls is described under their respective system.
IFF CONTROL PANEL CDFPX III

Refer to figure 177. The IFF control panel, located on
the left console, provides backup control of essential
CNI functions and some primary functions of IFF.
1185

T.O. GR1F16CJ1
D LOW - Same as NORM.

IFF Control Panel
C DF (Typical)

D NORM - Full range recognition and reply occur.
D EMER - The knob must be lifted to position to

EMER. When so positioned, an emergencyindicat
ing pulse group is transmitted each time a mode 1,
2, or 3/A interrogation is recognized. IFF replies tomode S interrogations with an alert consideration.

2

1

3

IFF M4 CODE Switch
The IFF M4 CODE switch has three positions. It is
springloaded to center (A/B) from HOLD and is leverlocked in ZERO position.

7

4
5

6

1. IFF MASTER Knob
2. IFF M-4 CODE Switch
3. C & I Knob
4. IFF MODE 4 MONITOR Switch
5. IFF MODE 4 REPLY Switch
6. IFF MODE 1/MODE 3 Selector Levers
7. IFF ENABLE Switch
GR1F-16CJ-1-2085X37

Figure 177.

C & I (Communications and IFF) Knob

Functions are:
D ZERO - Zeroizes mode 4 settings whenever IFF

MASTER knob is not OFF.

D A/B

- A/B position permits code selection via
upfront controls or via IFF MODE 4 REPLY switch
when C & I knob is in BACKUP.

D HOLD - Both code settings can be retained after

flight by placing the IFFM4 CODE switch to
HOLD momentarily prior to placing the IFF
MASTER knob to OFF or removing power.

IFF MODE 4 REPLY Switch

Functions are:

The IFF MODE 4 REPLY switch has three positions
and is used when the C & I knob is in BACKUP.

D BACKUP - In the event of failure of the upfront

Functions are:controls, the BACKUP position provides foralternate operation of the UHF and IFF. BACKUP
may be selected, when desired, even when upfront
controls are functioning.

D UFC - Provides for normal control of communica

tions, navigation, and IFF primarily via upfront
controls. Refer to UPFRONT CONTROLS, thissection.

IFF MASTER Knob
The IFF MASTER knob functions with the C & I knob
in BACKUP or UFC.
Functions are:
D OFF - Removes power from the IFF equipment andzeroizes mode 4 settings unless HOLD function has
been used. The knob must be lifted to position to orfrom OFF.

D STBY - The equipment is turned on and warmed up

but does not transmit.

1186

D OUT - Mode 4 operation is disabled.
D A - Enables mode 4 and selects the preset code for A.
D B - Enables mode 4 and selects the preset code for B.

IFF MODE 4 MONITOR Switch
The IFF MODE 4 MONITOR switch permits audio
monitor of mode 4 replies when the C & I knob is in
BACKUP.
Functions are:
D AUDIO-Monitoringof mode 4 interrogations isprovided by an audio tone on the intercom. An audio
tone of 0 to 1000 Hz is generated when thetransponder is not replying to valid mode 4interrogations. The frequency of the tone depends on
the number of interrogations received; i.e., the higher
the number of interrogations, the higher thefrequency.

D OUT - Disables audio monitor from intercom.

T.O. GR1F16CJ1
IFF ENABLE Switch
The IFF ENABLE switch permits mode selection in
BACKUP.

AUX COMM Panel
C DF (Typical)
1

Functions are:2

D M3/MS - Modes 3/A and S enabled in backup.

3

D OFF - Modes 1, 3/A, 4 and S disabled in backup.
D M1/M3 - Modes 1 and 3/A enabled in backup.

4

IFF MODE 1/MODE 3 Selector Levers
The MODE 1 code selector levers are incremented/de
cremented to select the twodigit code displayed in thereadout windows. The two leftmost windows display
the mode 1 code digits when the C & I knob is in
BACKUP.
The Mode 3/A code selector levers are incremented/
decremented to select the twodigit code displayed in
the readout windows. The two rightmost windows
display the two most significant mode 3/A backup
code digits when the C & I knob is in BACKUP. The
two least significant digits are always set to zero. For
example, if the windows are set to 77, the transmittedmode 3/A code is 7700.
AUX COMM PANEL CDF PX II

Refer to figure 178. The AUX COMM panel, locatedon the left console, provides backup control ofessential CNI functions and some primary functionsof IFF.
CNI (Communications, Navigation, and IFF)
Knob
Functions are:
D BACKUP - In the event of failure of the upfront

controls, the BACKUP position provides foralternate operation of the UHF, TACAN, and IFF(mode 4 and EMER functions only). BACKUP may
be selected, when desired, even when upfront
controls are functioning.

D UFC - Provides for normal control of communica

tions, navigation, and IFF primarily via upfront
controls. Refer to UPFRONT CONTROLS, thissection.

7

5

6

1. IFF MASTER Knob
2. M-4 CODE Switch
3. CNI Knob
4. TACAN Function Switch
5. TACAN CHANNEL Selector Levers
6. IFF MODE 4 MONITOR Switch
7. IFF MODE 4 REPLY Switch
GR1F-16CJ-1-0077X37

Figure 178.

IFF MASTER Knob
The IFF MASTER knob functions with the CNI knob
in BACKUP or UFC.
Functions are:
D OFF - Removes power from the IFF equipment andzeroizes mode 4 settings unless HOLD function has
been used. The knob must be pulled outward toposition to OFF from STBY.

D STBY - The equipment is turned on and warmed up

but does not transmit.

D LOW or NORM - Full range recognition and reply

occur. Transmitted power from the IFF system isthe same for both the LOW and NORM positions.

D EMER -The knob must be pulled outward toposition to EMER. When so positioned, an
emergencyindicating pulse group is transmitted
each time a mode 1, 2, 3/A, or S interrogation isrecognized.

M4 CODE Switch
The M4 CODE switch has three positions. It is
springloaded to center (A/B) from HOLD and is leverlocked in ZERO position.
1187

T.O. GR1F16CJ1Functions are:
D ZERO - Zeroizes mode 4 settings whenever IFF

MASTER knob is not OFF.

D A/B

- A/B position permits code selection via
upfront controls or via mode 4 REPLY switch whenin BACKUP.

D HOLD - Both code settings can be retained after

flight by placing the M4 CODE switch to HOLD
momentarily prior to placing the IFF MASTER
knob to OFF or removing power.

IFF MODE 4 REPLY Switch
The IFF MODE 4 REPLY switch has three positions
and is used when the CNI knob is in BACKUP
position.
Functions are:
D OUT - Mode 4 operation is disabled.
D A - Enables mode 4 and selects the preset code for A.
D B - Enables mode 4 and selects the preset code for B.

IFF MODE 4 MONITOR Switch
The IFF MODE 4 MONITOR switch permits audio
monitor of mode 4 replies when in BACKUP.
Functions are:
D AUDIO-Monitoring of mode 4 interrogations andreplies is provided by an audio tone on theintercom. An audio tone of 0 to 1000 Hz is
generated when the transponder is not replying tovalid mode 4 interrogations. The frequency of thetone depends on the number of interrogations
received; i.e., the higher the number of interrogations, the higher the frequency.

D OUT - Removes audio monitor from intercom.

TACAN Function Switch
The TACAN function switch provides control of
TACAN functions when the CNI knob is in BACKUP
position.
Functions are:
D T/R - Transmit/receive mode. Same as REC and in

addition, interrogates the TACAN ground station
for DME information; distance (nm) is displayedin the HSI range indicator (MILES).

1188

D REC - Receive mode. The system receives signals

which result in a bearing and course deviation
display on the HSI and audio in the headset.
TACAN range indicator (MILES) on HSI isshuttered.

D A/ATR - Airtoair transmitreceive mode.
TACAN system interrogates and receives signals
from aircraft having airtoair capability, provid
ing slant range (nm) distance between aircraft
operating 63 TACAN channels apart. (KC10A
also provides bearing information.) Up to fiveaircraft can determine distance from a sixth leadaircraft.

TACAN CHANNEL Selector Levers
Four TACAN CHANNEL selector levers andreadout windows are provided for TACAN opera
tion when the CNI knob is in BACKUP. The TACAN
channel is set into the first three windows. The
fourth window is used to select X or Y band.
AUDIO 1 CONTROL PANEL
Refer to figure 179. The AUDIO 1 control panel
provides control of less frequently used functions of
the communications system. Except as noted,
controls are active regardless of the position of the
CNI knob.
COMM 1 (UHF) Power Knob
The COMM 1 power knob has an OFF position anda cw arrow pointing to INC. The OFF position
removes power from the UHF radio. Rotating theknob cw applies power and increases UHF audio
volume.
COMM 1 (UHF) Mode Knob
The COMM 1 mode knob has three positions andmay be depressed in any of the three positions.
Depressing the knob interrupts reception andtransmits a tone signal and TOD for HQ on theselected frequency.
Functions are:
D OFF - Disables squelch circuit to permit receptionof a weak signal.

D SQL- Enables squelch circuit to help reduce
background noise in a normal operation.

D GD- The main receiver and transmitter areautomatically tuned to the guard frequency and theguard receiver is disabled. GD position is not
functional with CNI knob in BACKUP.

T.O. GR1F16CJ1

Audio Control Panels (Typical)
1
C

2

3

4

1

2

3

4

DF

DR

5
8

5
8

6

7

6

7
12

12
1. COMM 1 (UHF) Power Knob
2. COMM 2 (VHF) Power Knob
3. SECURE VOICE Knob
4. MSL Tone Knob
5. THREAT Tone Knob
6. TF Tone Knob

11

10

9

7. COMM 2 (VHF) Mode Knob
8. COMM 1 (UHF) Mode Knob
9. HOT MIC CIPHER Switch
10. ILS Power Knob
11. TACAN Power Knob
12. INTERCOM Knob

9

11

10

GR1F-16CJ-1-0078X37

Figure 179.

COMM 2 (VHF) Power Knob

D GD

The COMM 2 power knob has an OFF position and acw arrow pointing to INC. The OFF position removes
power from the VHF radio. Rotating the knob cw
applies power and increases VHF audio volume.

SECURE VOICE Knob

COMM 2 (VHF) Mode Knob

MSL Tone Knob

The COMM 2 mode knob has three positions and may
be depressed to transmit a VHF tone in any of thethree positions. The COMM 2 mode knob may not
function during certain failures of the upfront
controls.

The MSL tone knob has a cw arrow pointing to INC.
Rotating the knob cw increases the volume of the tone
from the AIM9 missile being monitored.

Functions are:

The TF tone knob is not functional.

D OFF - Disables squelch circuit to permit reception

THREAT Tone Knob

D SQLThe THREAT tone knob has a cw arrow pointing to
INC. CW rotation increases the volume of the TWS
composite tone.

of a weak signal.

- Enables squelch circuit to help reduce
background noise in normal operation.

- The main receiver and transmitter areautomatically tuned to the guard frequency.

The SECURE VOICE knob is inoperative.

TF Tone Knob

1189

T.O. GR1F16CJ1
AUDIO 2 CONTROL PANEL
Refer to figure 179. The AUDIO 2 control panel
provides control of less frequently used functions of
the communications system. Except as noted,
controls are active regardless of the position of the
CNI knob.

D OFF- Deactivates HOT MIC and CIPHER
functions.

D CIPHER - Momentary position which limits UHF

and VHF reception to secure voice only. CIPHER isfunctional only when operating in secure voicemode.

ILS Power Knob

ANT SEL PANEL C

The ILS power knob has an OFF position and a cw
arrow pointing to INC. The OFF position removes
power from the ILS receiver. Rotating the knob cw
applies power and increases the volume of thelocalizer identification signal. The ILS may not
function during certain failures of the upfront
controls.

Refer to figure 180. The ANT SEL panel allows
selection of various antennas for optimum transmis
sion and reception of IFF and UHF signals.

TACAN Power Knob
The TACAN power knob has an OFF position and a cw
arrow pointing to INC. The OFF position removes
power from the TACAN receiver. Rotating the knob
cw applies power and increases volume of the TACAN
station identification signal.
INTERCOM Knob
The INTERCOM knob has a cw arrow pointing to
INC. Rotating the knob increases the volume of thesignals available to the intercom set.
The intercom provides the following functions:DF

IFF ANT SEL Switch C

DF

The IFF ANT SEL switch is a threeposition switch.
Functions are:
D UPPER - Upper antenna is used to receive andreply to interrogation signals.

D NORM - The system selects the antenna which isreceiving the best signal.

D LOWER - Lower antenna is used to receive andreply to interrogation signals.

ANT SEL Panel
C DF (Typical)

D Monitoring and volume control of voice communica

tion between pilot and ground crew or between pilot
and tanker boom operator, monitoring and volume
control of AIM9 missile tone, TWS composite audio
tone, and TWS missile launch tone.

D Monitoring of systems individually volume con

trolled from the audio control panels.

D Monitoring of fixed volume warning tones (LG andlow speed warning tone, TWS missile launch tone,
and IFF mode 4 audio monitor) and voice messages.

HOT MIC CIPHER Switch
The HOT MIC CIPHER switch is a threeposition
switch.
Functions are:
D HOT MIC - Activates communication between pilot

and tanker boom operator or ground crew.
Activation of UHF VHF IFF transmit switch on thethrottle will override this function.

1190

1

2

1. IFF ANT SEL Switch
2. UHF ANT SEL Switch
GR1F-16CJ-1-0079X37

Figure 180.

T.O. GR1F16CJ1
UHF ANT SEL Switch C

DF

The UHF ANT SEL switch is a threeposition switch.
Functions are:
D UPPER - Upper antenna is used to receive andtransmit signals.

D NORM - The antennas cycle between upper andlower to provide omnidirectional antenna pattern.

D LOWER - Lower antenna is used to receive andtransmit signals.

COMMUNICATIONS SYSTEM
The aircraft communications system consists of a
HAVE QUICK UHF radio, a VHF radio, and anintercom set. Communications are controlled by the
upfront controls, two audio control panels, a UHFradio backup control panel, and an PX II AUX COMM
panel, PX III IFF control panel.
ANTENNA LOCATIONS

for channel 20. This tone indicates that the WOD
element for channel 20 was accepted and transferredto the radio. The WOD is stored in the nonvolatile
preset channel memory until a WOD transfer occurs.
This procedure overwrites preset channels 1520.
Single WOD Transfer
The WOD must be transferred from the nonvolatile
preset channel memory to the volatile WOD memoryfor HQ operation. Progressively select the next lower
channel (1915) until a double beep is heard atchannel 15. A double beep indicates that all WOD
elements have been transferred to the volatilememory.
When the HQ radio is off, the WOD data is not lost,
but is stored in nonvolatile preset memory. When theradio is initially turned on, the WOD must betransferred to the volatile memory. To transfer WOD
data to the volatile memory, select preset mode andrepeat the steps described above.
PX III Single WOD transfer is done automatically at
power up if the single WOD method was used forloading.

Refer to figure 181.
Multiple Word of Day (MWOD)
HAVE QUICK (HQ) SYSTEM
The HQ system provides normal and antijamming,
airtoair, and airtoground UHF communication
capability. The usual operating mode for an HQ UHF
radio is in the normal mode where the radio uses 1 of
7000 channels. The antijamming (AJ) mode uses afrequency hopping scheme to change the channel orfrequency many times per second. Because theparticular frequency used at any instant depends on
the precise time of day (TOD), both participating HQ
UHF radios must have clocks which are synchro
nized. In addition, the HQ UHF radio uses word ofday (WOD) and net number in the AJ mode.
Single Word of Day (WOD)
The WOD entry is normally entered before flight, but
it is possible to enter it in flight. WOD is entered byusing the six preset channels 1520. The WOD
elements are entered with the manual frequency
knobs and the PRESET button. For a new WOD
entry, start at preset channel 15 and enter an element
using the same method as in entering preset
frequencies in the normal mode. Progressively selectthe next higher preset channel (1620) and continue
to enter WOD elements into preset memory. A shorttone is heard when the PRESET button is depressed

MWOD capability allows for up to six WOD's anddates to be stored at one time into nonvolatile MWOD
memory. If the radio is turned off or power is lost after
entry of MWOD and date codes, the data is not lost;
therefore, the information remains intact until
manually changed. The six most recent WOD's are
retained. If a WOD with a duplicate date is entered,
the new entry takes precedence. After the MWOD
information is entered, the operational date must beentered so the radio can select the proper WOD. If
power is lost, the operational date is lost and must be
reentered. However, if a TOD is received, theoperational date is automatically entered as part of
the TOD message. To access MWOD load, selectpreset mode, enter 220.025 in channel 20, anddepress PX II PRESET, PX III LOAD (MLOAD isdisplayed) button. Next, select manual mode and load
word of day elements in channels 2015 by PX II
depressing the TONE button, PX III positioning theTTONE switch to TONE. After each entry is made,a tone is heard. After the first WOD is entered, thecorresponding date code is entered by selecting
channel 14. The date code is represented by theformat 3AB.000, where AB equals the day of themonth; e.g., 315.000 indicates the 15th of the month.
Press the PX II TONE button, PX III TTONE switch
to complete the date entry.

1191

T.O. GR1F16CJ1

Antenna Locations (Typical)
VHF AMVHF FM
UPPER IFF/UHF
ECM

THREAT WARNINGUPPER TACANGLOBAL POSITIONING SYSTEMIFF ARRAY

THREAT WARNINGECM

THREAT WARNINGAND ECMRADAR

ECM(RIGHT SIDE)ECM
RADAR ALTIMETERGLIDE SLOPE/LOCALIZERLOWER IFF/UHFIFF INTERROGATOR
MARKER BEACONLOWER TACANTHREAT WARNING
GR1F-16CJ-1-0080A37

Figure 181.

1192

T.O. GR1F16CJ1
A double beep indicates the date code has beenentered. When using MWOD, the current date must be
set into channel 1 prior to receiving TOD or the current
date must be part of the received TOD message.
Without date information, the radio cannot select thecurrent WOD, and a steady tone results when entering
AJ mode. After entering WOD(s) and/or setting
operational date, select 220.00, preset channel 20, anddepress PRESET button to enter the verify/operatemode. The UHF radio is inoperative when in any modeother than verify/operate. MWOD capability can be
determined by selecting 220.000, channel 20 presetmode, and pressing the PX II PRESET, PX III LOAD
(MLOAD is displayed) button. Select channel 19; if
one or two beeps are heard, the radio is not equippedwith MWOD capability. If no beep is heard, the radio
has MWOD capability. Verify a WOD for a particulardate is loaded by selecting channel 20, manual mode,and by selecting the date to be verified with themanual frequency knobs. Select CHAN 19, then
CHAN 20. A single beep signifies that a WOD for thatdate is present. The single WOD entry procedureremains valid for radios that are equipped for MWOD
operation. Preset channel 20 is reserved for MWOD
mode selection.
Automatic MWOD Loading PX III
The HAVE QUICK II PHASE 2 radio allows
automatic MWOD loading with a KYK13/TSEC filldevice. Insure radio power is on. Lift access door toreveal FILL connector and insure mode switch on filldevice is in OFF position. Install fill device in FILL
connector and place mode switch on fill device to ON.
Frequency/status display displays FILL. Selectdesired channel on fill device and depress LOAD
button. A series of beeps is heard and thefrequency/status display displays WOD OK if WOD isvalid and successfully received. If WOD is invalid or
unsuccessfully received, the frequency/status display
displays BAD. If BAD is displayed, the fill device
must be disconnected and reloaded and the loading
procedure must be repeated. If WOD OK is displayed,
repeat the loading procedure for subsequentchannels.
Insure WOD OK is displayed after each WOD isloaded. The CHAN display displays CHAN 1.
Operational date may be loaded by positioning thefrequency switches to OP DATE and setting theTTONE switch to TONE. After all WOD's are loaded,place mode switch on fill device to OFF anddisconnect FILL connector. Close access door. Radio
returns to previous mode and the frequency/status
display displays previous settings.

TimeofDay (TOD) Transmission
The TOD entry is normally entered before flight, but
it is possible to enter it in flight. It is possible totransmit timing information in both normal and AJ
modes by momentarily pressing the PX II TONE
button, PX III TTONE switch. In the normal mode, acomplete TOD message is transmitted, while in the
AJ mode, only an abbreviated time update istransmitted. A time transmission allows a time
update if one radio has drifted out of synchronization.
TimeofDay (TOD) Reception PX III
The HAVE QUICK radio accepts the first TOD that isreceived after power up. When the EGI is first poweredon, the EGI TOD signal will likely not be synchronizedwith the GPS constellation universal time coordinated(UTC) time since 3045 seconds are typically required
for satellite acquisition. To insure that the HAVEQUICK radio is using a synchronized UTC TOD,
another TOD should be manually received after anadequate GPS track has been achieved. The bestindication of an adequate GPS track is to access the
NAV STATUS DED page (LIST 4) and verify thehorizontal accuracy status. A GPS HIGH estimatedhorizontal accuracy status (or 300 feet if selected
accuracy display format is in feet) should insure that
the EGI TOD signal is synchronized with UTC time.
Alternatively, verify that four satellites are being
tracked by the EGI by confirming that a 4 is displayedfor EGI miscellaneous parameter (EGIM) 200.
TimeofDay (TOD) Reception PX II
Reception is possible in both normal and AJ modes.
The radio automatically accepts the first TOD
message received after the radio is turned on.
Subsequent messages are ignored unless the T
position is selected with the A32T knob. The radio
then accepts the next TOD update in either normal or
AJ mode, provided the TOD update arrives within 1
minute of the time the T position has been selected.
To receive time in the normal mode, position theA32T knob to T and return to a normal channel ineither manual or preset mode. To receive a time
update in AJ mode, rotate the A32T knob to the T
position and then back to the A position. A TOD
update (time tick) can now be received on the selected
AJ net for 1 minute.

Change 11193

T.O. GR1F16CJ1Net Number
After TOD and WOD are entered, any valid AJ net
number can be selected by using the manualfrequency knobs.
Antijamming (AJ) Mode Operation
An intermittent tone is heard in the headset if aninvalid AJ net is selected. A steady tone is heard if
TOD was not received or if a WOD for the current date
was not entered.
In the AJ mode, the radio can receive and process two
simultaneous transmissions on the same net.
Conferencing is controlled by the WOD loaded inchannel 19. If the WOD in channel 19 ends with 00 or
50, conferencing is possible. If the WOD in channel 19
ends with 25 or 75, conferencing is disabled.
If the function knob is set to BOTH and guard channel
jamming is encountered, switching the function knob
to MAIN negates the jamming.
HAVE QUICK (HQ) UHF RADIOThe UHF radio provides lineofsight communica
tions. UHF transmissions are made by holding the
UHF VHF IFF transmit switch on the throttle to theUHFposition.
Frequencies
range
from
225.000399.975 MHz. The guard receiver monitors
the guard frequency of 243.0 MHz. Power, volume,
squelch, guard, and tone controls are located on
AUDIO 1 control panel. HAVE QUICK mode selection
and other UHF functions are controlled by the
upfront controls.
Twenty or 19 channels may be preset using upfront or
backup controls, respectively. The UHF radio is
powered by battery bus No. 1 and requires power fromemergency ac bus No. 2 and emergency dc bus No. 2
for operation on upfront control.
Operation of the UHF Radio on Upfront Controls
To change UHF preset channels on the CNI page,refer to CNI READOUT/DED PAGE SUMMARY, thissection.
The UHF page is selected by depressing the COM 1
override button on the upfront controls.

1194Change 1

Functions available on the UHF page are:
D Changing preset channels and manual frequencies.

D Changing between MAIN and BOTH functions.
D Channelizing frequencies.
D Changing bandwidth.
D Updating TOD and selecting AJ mode.

The UHF page displays its data on five lines. The firstline contains information about the radio mode andtransmission status. The radio mode displays
UHFTOD for normal operation; HQTOD isdisplayed when the AJ mode is selected. Either label
is highlighted when the radio is keyed. The
transmission status is displayed as OFF, MAIN, or
BOTH. OFF is displayed when power is turned off
with the COMM 1 power knob.
PX III Following UFC power cycles with weighton
wheels, BOTH is automatically displayed; whenairborne, the previously selected mode (MAIN or
BOTH) is retained.

MAIN is displayed when transmissions are received
only on the main UHF receiver; BOTH is displayedwhen transmissions are received on both the mainand guard receivers. The second line displays theactive frequency or preset channel number. Line
three displays the data entry scratchpad. The fourthline displays the label PRE and a preset channel
number, increment and decrement symbols, and thelabel TOD.
Line five displays the frequency associated with thepreset channel and selected bandwidth, which isdisplayed as either narrow band (NB) or wide band
(WB). PX III At UFC power up and weightonwheels,
NB is automatically selected. PX III The CNI
readout/DED page summary shows the radio mode
UHF, transmission status (BOTH), active presetchannel number (10), the scratchpad frequency(*235.00*), a preset channel (PRE 10), and label(TOD), the frequency associated with the preset
number in line four (226.00), and radio bandwidth(NB). PX II The CNI readout/DED page summary
shows the radio mode (UHFTOD), transmission
status (MAIN), active preset channel number (10),the scratchpad frequency (*235.00*), a presetchannel (PRE 10), the frequency associated with thepreset channel in line four (226.00), and radio
bandwidth (NB).

T.O. GR1F16CJ1Changes on the UHF page are made by the use ofasterisks and the keyboard, the increment/decrement
switch, or the SEQ position of the DCS, depending on
the nature of the change. Asterisks (*235.00*) are
moved by up/down movement of the DCS through thefollowing locations: scratchpad, TOD, bandwidth(UHF mode only), radio mode, preset channel, presetchannel frequency designation, and back to thescratchpad. Information enclosed by the asterisks
may be changed. With asterisks properly positioned,
the radio mode and bandwidth are changed bydepressing any ICP/IKP key (19). Channels/frequen
cies are changed by placing the asterisks over thescratchpad, keying in the desired channel/frequency,
and then depressing the ENTR button. If one or two
digits are entered, a preset channel is selected; if
three to five digits are entered, a frequency is
selected. Extra numbers over five digits are ignored.
When the ENTR button is depressed, the channel/fre
quency changes and the DED reverts to theoverridden page. (If the COM 1 override button isdepressed and the ENTR button is depressed without
keying new digits in the scratchpad, the channel orfrequency that was previously in the scratchpad isselected.) The preset channel is changed by using
either the increment/decrement switch (note theup/down triangles next to the preset channel number)or by use of the asterisks and keyboard as previously
described. As the preset channel is changed, its
associated frequency also changes. Changing
between MAIN and BOTH functions can only be
accomplished by positioning the DCS to SEQ. Thisaction can be accomplished independently of asterisk
position. For HAVE QUICK selection, enter 0 in thescratchpad.
HAVE QUICK primary page displays data on fivelines. The first line displays HQ radio mode with
submode characters (HQTNG or HQCBT). PX III
The receiver mode (BOTH, MAIN) is displayed on
the first line and can be changed by positioning the
DCS to SEQ. PX II The net types for training
submode are displayed TNET or FMT; combat
submode net types are A/B, NATO, or NNTO. Linetwo displays PX III the active net number/preset
and the net type, PX II the active net number/pre
set. When using combat HAVE QUICK frequencies,
the first three digits indicate the HAVE QUICK net
number. The last two digits indicate the combat
submode where 00 corresponds to A/B, 25 to NATO,and 50 to NNTO. For example, A32.425 is combat
net 324, submode NATO.

Line three displays data entry scratchpad, where newactive net numbers or preset numbers are keyed inand entered. Changing the active preset via thescratchpad also sets the preset number to the samevalue. Line four displays PRE mnemonic, presetchannel number and TOD label. The system supports
entry/storage of 20 net numbers in preset locations.
The fifth line displays the preset net number.
Asterisks are moved by up/down movement of the
DCS through the following locations: scratchpad,
TOD, HQ submode, preset channel number, preset
net, and back to the scratchpad.
Special considerations should be given in instanceswhere the features of the HAVE QUICK radio areused. The HAVE QUICK radio operates in twomodes:UHF, for the normal mode using a single
UHF frequency, and HQ, for the AJ mode using
pseudorandom
frequencyhopping
techniques.
Before the AJ mode can be used a WOD must be
loaded, a TOD must be received, and a valid net must
be selected. The radio automatically accepts the first
TOD message received after the radio is turned on.
After the initial TOD is received an updated TOD canbe received by placing the asterisks around the TODlabel on the UHF page and then depressing the ENTR
or MSEL button. The asterisks autostep back to thescratchpad and TOD is highlighted. A TOD updatemust be received within 1 minute of depressing the
ENTR or MSEL button. TOD label dehighlights after
1 minute or when a scratchpad operation is
performed. If the TOD label is highlighted when an
UFC power cycle or an autorestart occurs TOD label
is dehighlighted upon power up. If the TOD label ishighlighted and the COM 1 HQ page is exited (for areason other than a scratchpad autoexit operation),
the TOD label remains highlighted on the COM 1 HQ
page (if the page is reentered before the 1 minute timedelay has timed out). After the TOD update and withproper WOD entries on the UHF backup control
panel, the AJ mode is accessed by entering zero on thescratchpad which gives the ability to switch between
HQ and UHF modes. There is no autoexit when zerois entered on the scratchpad. When HQ is selected,
the bandwidth window is blanked from the page.
Nets, consisting of three digits, are entered andchanged just as operating frequencies and can bepreentered in preset locations.

Change 11195

T.O. GR1F16CJ1Mode Knob PX II

UHF Radio Backup Control Panel
Refer to figure 182. The UHF radio backup control
panel, located on the left console, controls the UHFradio, when required, by positioning the PX II CNI
knob to BACKUP (AUX COMM panel), PX III C & I
knob to BACKUP (IFF control panel. When BACKUP
is selected, the controls on the UHF radio backup
control panel have the following functions:

Functions are:
D MANUAL - UHF frequency is selected by manuallysetting the five frequency knobs.

D PRESET - UHF frequency is determined by the

CHAN knob.

D GUARD - The main receiver and transmitter areautomatically tuned to the guard frequency and theguard receiver is disabled.

Function Knob
Functions are:Mode Knob PX III
Functions are:

D OFF - Power off.
D MAIN - With COMM 1 power switch on, UHF radio

operates on selected frequency.

D MNL - UHF frequency is selected by manuallysetting the five frequency knobs.

D PRESET - UHF frequency is determined by the

D BOTH - Normal operation plus receiving on guard

frequency.

CHAN knob.

D GRD- The main receiver and transmitter areautomatically tuned to the guard frequency and theguard receiver is disabled.

D ADF - Not operational.

UHF Radio Backup Control Panel C DF
1 2
(Typical)
3 4 5
18
17
16

6

14

7

5

18

7

15

13

12

10

9

8

13

10

9

8

PX III

PX II
1. ZERO Switch
2. FILL Connector
3. CHAN Display
4. Frequency/Status Display
5. CHAN Knob
6. STATUS Button

11

7. Manual Frequency Knobs
8. Mode Knob
9. SQUELCH Switch
10. VOL Knob (Inactive)
11. T-TONE Switch
12. TONE Button

13. Function Knob
14. A-3-2 Knob
15. A-3-2-T Knob
16. TEST DISPLAY Button
17. LOAD Button
18. Preset Channel Card and Access Door
GR1F-16CJ-1-0081A37

Figure 182.

1196

T.O. GR1F16CJ1
CHAN Knob

VOL Knob

The CHAN knob permits the selection of 1 of 19(MWOD) or 14 (single WOD) preset frequencies(channels 119) with the mode knob at PRESET and
the PX II A32T, PX III A32 knob in either 2 or 3.
Preset channels used for WOD storage cannot be usedas preset channels for normal radio operation.
Frequencies set for each channel can be manually
written on a channel frequency card located on theaccess door. Preset channel frequencies are set (orchanged) as follows:

The VOL knob is nonfunctional. Volume can only becontrolled by the COMM 1 (UHF) power knob.

D Function knob - MAIN or BOTH.
D Mode knob - PRESET.
D Manual frequency knobs - Set to desired frequency.
D CHAN knob - Set to desired channel.
D Lift access door.
D Depress PX II PRESET, PX III LOAD button underaccess door.

SQUELCH Switch
Functions are:
D ON

- Enables squelch circuit which helps toeliminate background noise in normal reception.

D OFF - Disables squelch circuit to permit receptionof a weak signal.

TONE Button PX II
Depressing the TONE button in normal or AJ modeinterrupts reception and transmits a tone signal and
TOD for HQ on the selected frequency. Simultaneous
ly pressing the TONE button in conjunction with theA32T knob in T position starts the emergency
startup of the TOD clock. Pressing the TONE button
with channels 1, 14, 1520 selected in manual mode
enters MWOD data when in load mode. Pressing the
TONE button in erase mode erases all MWOD data.
CHAN Display PX III

Manual Frequency Knobs
The five manual frequency knobs allow manual
selection of frequencies in steps of 0.025 MHz from
225.000399.975 MHz.

The CHAN display displays selected channel number
when mode knob is positioned to PRESET. If in
MLOAD or FMT.CNG operating mode, the selected
memory location is displayed. Display is blank whenthe mode knob is positioned to MANUAL or GUARD.

A32T Knob PX II

Frequency/Status Display PX III

Functions are:

The frequency/status display displays several condi
tions of radio operation. Four programming modes
control AJ operation. These modes are VER/OP,
MLOAD, ERASE, and FMT.CHG. These modes areaccessed by placing the CHAN knob to 20, rotating
the mode knob to PRESET, and keying in theappropriate frequency for the applicable mode.
Frequencies are 220.000 for VER/OP, 220.025 for M
LOAD, 220.050 for ERASE, and 220.075 for
FMT.CHG. VER/OP is displayed when the STATUS
button is depressed. MLOAD is displayed when
MWOD's are loaded manually. ERASE is displayedwhen WOD's are erased by depressing the TONE
button or by selecting 0. FMT.CHG is displayed whenfrequency management training net frequencies arechanged by selecting a channel with the mode knob
and changing frequencies with the manual frequency
knobs.

D A - Selects AJ mode.
D 3 - Allows manual selection of frequencies.
D 2 - Allows manual selection of frequencies.
D T - Momentary position which enables the radio toaccept a new TOD for up to 1 minute after selection.
Also used in conjunction with the TONE button for
emergency startup of the TOD clock when TOD is
not available from external sources.

A32 Knob PX III
Functions are:
D A - Selects AJ mode.
D 3 - Allows manual selection of frequencies.
D 2 - Allows manual selection of frequencies.

Other displays are:
D FILL- A fill device is connected to the FILL
connector to load MWOD's.
Change 11197

T.O. GR1F16CJ1
D BAD - A WOD was unsuccessfully received.
D WOD OK - A correct WOD has been successfullyreceived.

STATUS Button PX III
When the STATUS button is depressed, selected
preset channels and manual frequencies may beverified on proper displays.

all normal operations. Power, volume, squelch, guard,
and tone controls for the VHF radio are located on the
AUDIO 1 control panel. Other VHF radio functions
are controlled by the upfront controls. There are no
backup controls. In the event of certain failures of the
upfront controls, the VHF radio may remain on thelast frequency set prior to the failure. The VHF radiois powered by PX III emergency dc bus no. 1, PX II thenonessential dc bus.

TTONE Switch PX III

Operation of the VHF Radio on Upfront Controls

Functions are:

To change VHF preset channels on the CNI page,refer to CNI READOUT/DED PAGE SUMMARY, thissection.

D T - Momentary position which enables the radio toaccept a new TOD for up to 1 minute after selection.
Also used in conjunction with the TEST DISPLAY
button for emergency startup of the TOD clock
when TOD is not available from external sources.

D TONE - In normal or AJ mode, the TONE positioninterrupts reception and transmits a tone signal
and TOD for HQ on the selected frequency.
Selecting TONE with channels 1, 14, 1520 selected
in manual mode enters MWOD data when in LOAD
mode. Selecting TONE when in ERASE modeerases all MWOD data.

TEST DISPLAY Button PX III
Depressing the TEST DISPLAY button illuminates
all segments of the frequency/status display and theCHAN display for a functional test.
ZERO Switch PX III
Depressing the ZERO switch erases all MWOD data
regardless of radio switch/knob positions or radio
operating mode.
FILL Connector PX III
The FILL connector connects fill device to the radio.
The FILL connector allows automatic MWOD
loading.
VHF RADIOThe VHF radio provides lineofsight voice commu
nication. VHF transmissions are made by holding the
UHF VHF IFF transmit switch on the throttle to the
VHF position. Transmission and reception areavailable for AM from 116.000151.975 MHz and for
FM from 30.00087.975 MHz. Only reception isavailable from 108.000115.975 MHz. Twenty channels may be preset. Operation may be either onnarrow band or wide band. Narrow band is used for1198

The VHF page is selected by depressing the COM 2
override button on the upfront controls. Functions
available on the VHF page are:
D Changing preset channels and manual frequencies.
D Channelizing frequencies.
D Changing bandwidth.

The VHF page displays its data on five lines. The firstline displays information about secure voice, the VHF
radio mode and status. The radio mode always
displays VHF for normal operation. The radio status
is displayed as ON or OFF; however, the ON and OFF
functions can only be controlled via the AUDIO 1
control panel. The second line displays active
frequency or preset channel. Line three displays thedata entry scratchpad, which is a fivedigit area
highlighted by two asterisks. The scratchpad is used
to change either channels or manual or active
frequencies, and it displays the channel/frequency
not currently being used. When the VHF page is first
selected, the scratchpad is displayed in one of two
ways:(1) If the radio is operating on manualfrequency, the scratchpad displays preset channel
entry. (2) If the radio is operating on preset channels,
the scratchpad displays manual or active frequency
entry. The fourth line contains the label PRE and apreset channel number. The label PRE is highlighted
when the radio is tuned to a preset channel. Line
five displays the frequency associated with thepreset channel. The fifth line also shows theselected bandwidth, which is displayed as either
narrow band (NB) or wide band (WB). The CNI
readout/DED page summary displays the radio
mode (VHF), radio power status (ON), active presetchannel number (12), the scratchpad frequency(*135.00*), a preset channel (PRE 12), thefrequency associated with the preset channel in line
four (126.00), and bandwidth (NB).

T.O. GR1F16CJ1Changes on the VHF page are made by use of theasterisks and the keyboard or the increment/decrement switch, depending on the change. Asterisks(*135.00*) are moved by up/down movement of the
DCS through the following locations: scratchpad,
bandwidth, manual frequency, preset channel
number, preset channel frequency designation, andback to the scratchpad. Information enclosed withasterisks may be changed as described under
OPERATION OF THE UHF RADIO ON UPFRONTCONTROLS, this section. The preset channel may
be changed by using either the increment/decrement switch (note the up/down triangles next to thepreset channel number) or by use of the asterisks
and the keyboard.
When operating on guard frequency, AM or FM
selections are available on the VHF page. To switchfrom AM to FM or FM to AM, depress any ICP/IKP
key (19).

Seat ejection also automatically performs an escapezeroize operation; refer to EJECTION SEATOPERATION, this section.

NAVIGATION SYSTEMSEMBEDDED
GLOBALPOSITIONING
INERTIAL NAVIGATION SET (EGI) PX III

AND

The EGI is the prime sensor for aircraft velocity,
attitude, heading, position, and is a source ofnavigation information. The EGI consists of anembedded GPS receiver (EGR) and ring laser gyro
inertial navigation unit.

ZEROIZE SWITCHThe EGR is a 12 channel receiver for allinview
tracking. The EGI provides GPSonly and freeinertial navigation solutions.

Refer to figure 13. The ZEROIZE switch is a guarded
threeposition toggle switch.

The EGI, in conjunction with the UFC, CADC, andMMC, provides:

When positioned to OFP, coded electronic information
associated with the following equipment is purged:

D Acceleration.

D DTC.

D Velocity.

D PX II FCC.
D PX III MMC.
D GPS.
D AIFF mode 4.
D PX III IDM.

D Position.
D Attitude.
D Magnetic and true heading.
D Altitude.

D PDG.
D RWR.

When positioned to DATA, coded electronic
information associated with the following equipment
is purged:
D DTC.
D GPS.
D AIFF mode 4.
D PX III MMC.
D PX III IDM.

D Angular rates.
D Universal time coordinated (UTC) and time tags.
D Instantaneous and maximum g data for display in

the HUD.

The EGI generates signals which drive the aircraft
attitude direction indicator (ADI) and horizontal
situation indicator (HSI).
For a detailed system description, refer to T.O.
GR1F16CJ3411.
1199

T.O. GR1F16CJ1GLOBAL POSITIONING SYSTEM (GPS) PX IIThe GPS receives signals from orbiting satellites todetermine accurate aircraft position, velocity, andtime information. The FCC uses this data to reduce
inertial navigation errors and enable accurate
weapon delivery solutions. GPS data can also be
used for inflight alignment of the INS. The FLCS
uses GPS data for SWIM during TF. GPS data is also
used in HAVE QUICK timeofday updates.
Refer to T.O. GR1F16CJ3411 for a detailed
description of GPS.
INERTIAL NAVIGATION SET (INS) PX IIThe INS is a prime sensor for aircraft velocity,
attitude, and heading and is a source of navigation
information.
The INS, in conjunction with the upfront controls,
GPS, CADC, and FCC provides:
D Present position with update and storage capability.
D Current winds.
D Groundspeed and drift angle.
D Greatcircle course computation with steering
provided to 30 points.

Operation of the TACAN on Upfront Controls
The TACAN is controlled by the upfront controls
when the TACAN power knob on the AUDIO 2
control panel is placed to on (cw) and the PX II CNI,PX III C & I knob is in UFC. The TCN/ILS page isselected by verifying the CNI page is displayed on
the DED and then depressing the TILS function
button on the upfront controls. The DED displays
information about three TACAN items: operating
mode, channel number, and band. The operating
mode is displayed (REC, T/R, or A/A TR) and may bechanged by positioning the DCS to SEQ until thedesired mode is displayed. Channel (1126) andband (X or Y) may be changed within the scratchpadwith the keyboard.
Channels are selected by keying in the desired
number on the scratchpad and then depressing the
ENTR button. Asterisks remain about thescratchpad after channel change. TACAN band isselected by keying zero in the scratchpad and thendepressing the ENTR button. The asterisks remain
about the scratchpad after band change.

TERRAIN FOLLOWING RADAR (TFR)
SYSTEM
The TFR system provides manual TF and ATF
capability. In manual TF, the NVP provides pitch
steering cues via the HUD. In ATF, climb/divecommands are calculated in the NVP and sent to the
FLCS to maintain a selectable clearance above theterrain.

D Instantaneous and maximum g data for display in

Refer to T.O. GR1F16CJ3411 for expanded TFRsystem description, operating procedures, andoperating limitations.

For a detailed system description, refer to T.O.
GR1F16CJ3411.

AIFF SYSTEMTACTICAL AIR NAVIGATION (TACAN) SYSTEM

The AIFF system provides selective identification
feature (SIF), automatic altitude reporting, andencrypted mode 4 IFF. Normal operation is possible
in any of six modes:

the HUD.

The TACAN system provides continuous bearing
and distance information from any selected TACAN
station within a lineofsight distance up to
approximately 390 miles, depending upon terrain
and aircraft altitude. Only distance information ispresented when a DME navigational aid is selected.
There are 252 channels available for selection. Twoantennas, one on top and one on the bottom of thefuselage, provide omnidirectional coverageregardless of aircraft attitude. The TACAN bearing,
selected course, range, and course deviation
information are displayed on the HSI as determinedby HSI and INSTR MODE knob settings.
1200

D Mode 1-Security identity.
D Mode 2-Personal identity.
D Mode 3/A-Traffic identity.
D Mode 4-Encrypted identity.
D Mode C-Altitude reporting.
D Mode S- PX II Data Link (Surveillance mode),
PX III Air traffic control data link (includes mode

3/A and C functions).

T.O. GR1F16CJ1AIFF TRANSPONDER PX II
The transponder only transmits coded replies tocorrectly coded interrogations. Backup control isprovided for mode 4 and EMER functions. Refer to
AUX COMM PANEL, this section, for functions of
IFF MASTER knob and mode 4 backup controls. In
the event of certain failures of upfront controls, IFF
modes 1 and 3/A may continue to reply on the lastcodes selected prior to failure.
Modes 2 and C, if selected prior to failure, may
continue to reply. However, when the CNI knob isplaced in BACKUP, modes 1, 2, 3/A, C, and S are
disabled except when the IFF MASTER knob is in
EMER. Mode C provides altitude information fromthe CADC in 100foot increments. The AIFF
transponder also provides for selection of manual or
automatic operation. If the automatic function is
selected, IFF operation can be programmed to:
D Turn modes on or off when a certain position isreached.

D Change codes when a specific time is reached.
D Turn modes on or off when a certain position isreached and changes codes when specific time isreached.

If the automatic position switching mode isselected, two mode groups are available. Modegroups contain selection status and position
criteria that determine where a selected group
becomes active. Mode groups are designated 1 or 2.
Mode groups may be preprogrammed via the DTE or
manually.
If the automatic time switching mode is selected, sixcode groups for IFF modes 1, 3/A, 4, and S areavailable. Code groups (16) designate the codes forthe four IFF modes and a time when the selectedcode group becomes active. Code groups may be
preprogrammed via the DTE or manually.
AIFF TRANSPONDER PX III
The transponder only transmits coded replies tocorrectly coded interrogations. Backup control isprovided for modes 1, 3/A, S, and 4 and EMER
functions. Refer to IFF CONTROL PANEL, thissection, for functions of IFF MASTER knob andmodes 1, 3/A, S, and 4 backup controls. In the eventof certain failures of upfront controls, IFF modes 1, 2,
and 3/A may continue to reply on the last codes
selected, if modes are selected prior to failure. ModeC, if selected prior to failure, may continue to reply.
When the C&I knob is placed in BACKUP, modes 2
and C are disabled. Modes 1, 3/A, 4, and S can be
enabled via switches on the IFF control panel. When
the C&I knob is placed in UFC, all modes areautomatically enabled. Mode C provides altitude
information from the CADC in 100foot increments.
The AIFF transponder also provides for selection ofmanual or automatic operation. If the automaticfunction is selected, IFF operation can be
programmed to:
D Turn modes on or off when a certain position isreached.

D Change codes when a specific time is reached.
D Turn modes on or off when a certain position isreached and change codes when a specific time isreached.

IFF modes may be selected and programmed for use
in groups. Mode groups contain selection status andposition criteria that determine where a selectedgroup becomes active. Mode groups are designated 1
or 2. Mode groups may be preprogrammed via theDTE or manually by placing the DED asterisks
around the scratchpad and entering the desired datausing the ICP/IKP keys.
If the automatic time switching mode is selected, sixcode groups for IFF modes 1, 3/A, and 4 are available.
Code groups (16) designate the codes for the three
IFF modes and a time when the selected code group
becomes active. Code groups may be preprogrammed
via the DTE or manually by placing the DED
asterisks around the scratchpad and entering thedesired data using the ICP/IKP keys.
Operation of the Transponder on Upfront
Controls PX II
When the IFF MASTER knob, located on the AUX
COMM panel, is placed to LOW or NORM and the
CNI knob is in UFC, the AIFF transponder iscontrolled by the upfront controls. The IFF page isselected by depressing the IFF override button on
the upfront controls. The IFF page displayed
depends on the mode/code switching function last
selected; i.e., the MAN page is displayed if manualfunction was last selected. IFF pages are accessedby placing the DCS to SEQ. Information isdisplayed on the IFF pages in four lines (lines one,
three, four, and five). Refer to CNI READOUT/DED
PAGE SUMMARY, this section.

1201

T.O. GR1F16CJ1
The first line displays IFF and power status (OFF,STBY, or ON). The next area displays MAN for
manual IFF operation or AUTO for automatic IFF
operation. If POS page is selected, the label POS isdisplayed in the next area of line one. The next area
displays the selected mode group (1 or 2). The area atthe far right displays two items: the label TIM isdisplayed if TIM is selected; the selected code group(16) for the IFF modes 1, 3/A, 4 and S is displayednext to the TIM label.
Line three displays four items. M1 is displayed at the
far left. Displayed with the M1 label is the twodigit
mode 1 code. The next area displays the MC label anda 5. MC represents IFF mode C, and it is highlighted
when mode C is selected (MAN and POS only). The 5
represents the ICP/IKP key that provides selection ofmode C if the ICP/IKP key is entered by thescratchpad. The area at the far right displays thescratchpad, which is displayed only when values arebeing keyed in. The scratchpad provides on/off control
for each of the six IFF modes, code control for modes
1, 2, 3/A, 4 and S, and monitoring control for mode 4.
Line four displays five items. The first area at the left
displays the M2 label, representing mode 2, andenterable mode 2 code (MAN only). The second area
displays M4 and an A or a B. M4 represents mode 4.
A or B indicates the two mode 4 codes. The third area
displays a 6, which represents the ICP/IKP key thatalternates selection of the mode 4 code (A or B). The
fourth area displays the criteria used for position
switching or time switching. This area displays adirection (N, S, E, or W), the fixed label OF, and a
destination number (125) (POS only) or a time
(HH:MM) (TIM only).
Line five displays data concerning IFF modes 3/A and
4 (MAN only) and S. The first area displays the label
M3 and the enterable fourdigit mode 3/A code. The
second area displays mode 4 monitoring modes (OUT,
LIT, AUTO). The third area displays a 7, which
represents the ICP key that rotates selection of mode
4 monitoring (MAN only). The fourth area displays
the label MS and an 8, which represents the ICP key
that rotates selection of mode S.
IFF manual mode switching is accomplished byverifying the IFF MAN page. With asterisks
positioned about the scratchpad, depress appropriate
ICP/IKP key and depress ENTR. ICP/IKP keys
represent the following:0, 8, and 9 are invalid(scratchpad flashes); 1 for mode 1 selection; 2 formode 2; 3 for mode 3/A; 4 for mode 4; 5 for mode C; 6
for mode 4 A/B code; 7 for mode 4 monitoring (OUT,

1202

LIT, or AUD); 8 for mode S selection; two digits (mode
1 code); 2 plus four digits (mode 2 code); three or four
digits (mode 3/A code).
IFF automatic switching is accomplished on the IFF
AUTO pages. IFF AUTO pages are accessed byplacing DCS to SEQ. To select AUTO POS or AUTO
TIM switching functions, position asterisks about the
POS and/or the TIM labels and depress the MSEL
button. If AUTO POS switching function is selected,
only position and mode criteria is displayed. Labels(6) (mode 4 codes) and (7) (mode 4 audio monitor
function) are removed and mode 4 A/B codes are notmodifiable. The scratchpad is reduced to one digit(mode numbers only). The mode group number isdisplayed on the page and is highlighted when thatmode group is active. The DCS provides access to thedifferent mode groups by positioning the switch to
INC or DEC. Sequencing to a different mode groupdoes not affect the active squawks.
If AUTO TIM switching function is selected, only timeand code criteria is displayed. Labels (5) (mode C is
not modifiable from this page) and (7) (mode 4 audio
monitor), and (8) (mode S) are removed. Mode 2 codes
are not displayed (only modifiable from the IFF MAN
page). The code group number is displayed on thepage and is highlighted when that code group is
active. The DCS provides access to the different codegroups by positioning the switch to INC or DEC.
Sequencing to a different code group does not affectthe active squawks.
If the FCC fails, only IFF manual mode is available.
Operation of the Transponder on Upfront
Controls PX III
Refer to figure 177. When the IFF MASTER knob,
located on the IFF control panel, is placed to LOW orNORM and the C&I knob is in UFC, the AIFF
transponder is controlled by the upfront controls. The
IFF page is selected by depressing the IFF override
button on the upfront controls. Information isdisplayed on the IFF pages in five lines.
The first line displays IFF and power status (OFF,STBY, DEGR, or ON). The next area displays MAN
for manual IFF operation or AUTO for automatic IFF
operation. The third area displays POS for position
switching and a two state mode group or TIM for timeswitching and a six state Code group.
Line two is blank.

T.O. GR1F16CJ1Line three displays four items. The first area at theleft displays M1 and the enterable mode 1 code (MANand TIM pages only). The second area displays M4
and an A or a B (MAN and TIM pages only). M4
indicates that mode 4 is selected. A or B indicates thetwo mode 4 codes. The third area displays a 6, which
represents the ICP/IKP key that alternates selectionof the mode 4 code (A or B). The area at the far rightdisplays the scratchpad, which is displayed only
when values are being keyed in, provides on/off foreach of the six IFF modes, code control for modes 1, 2,
3/A, and 4, and monitoring control for mode 4.
Line four displays five items. The first area at the left
displays M2 and the enterable mode 2 code (MANpage only). The next area displays MC and 5. MC
represents IFF mode C, and it is highlighted whenmode C is selected (MAN and POS pages only). The 5
represents the ICP/IKP key that provides selection ofmode C via the scratchpad. The fourth area displays
the criteria used for position switching. This area
displays a direction (N, S, E, or W), the fixed label OF,
and a destination number (125) (POS page only).
Line five displays five items. The first area displays
M3 and the enterable mode 3/A code (MAN and TIM
pages only). The second area displays mode 4
monitoring modes (MAN page only). The third area
displays a 7, which represents the ICP/IKP key thatrotates selection of mode 4 monitoring. The fourtharea displays MS. The fifth area displays an 8, which
represents the ICP/IKP key that provides selection ofmode S via the scratchpad.
The initial IFF page display is always MAN. IFF
pages are accessed by placing the DCS to SEQ. IFF
manual mode switching is accomplished by verifying
the IFF MAN page is selected. With asterisks
positioned about the scratchpad, depress appropriate
ICP/IKP key and depress ENTR. ICP/IKP keys
represent the following:0 and 9 are invalid(scratchpad flashes); 1 for mode 1 selection; 2 formode 2; 3 for mode 3/A; 4 for mode 4; 5 for mode C; 6
for mode 4 A/B code; 7 for mode 4 monitoring (OUT,
LIT, or AUD); 8 for mode S selection; any two digits(mode 1 code); three or four digits (mode 3/A code); 2
followed by 4 digits (mode 2 code). IFF automatic
switching is accomplished on the POS and TIM pages.
These pages are accessed by placing DCS to SEQ. To
select AUTO POS or AUTO TIM switching functions,place DCS to SEQ to display the POS and TIM pages.

If AUTO POS switching function is selected, only
position and mode criteria are displayed.
The scratchpad is reduced to one digit (mode numbersonly). The active mode group is displayed on the pageand is highlighted when that code is active. The DCS
provides access to the different mode groups bypositioning the switch to INC or DEC. Sequencing toa different mode group does not affect the active
squawks. If AUTO TIM switching function is
selected, only time and code criteria are displayed.
The active code group is displayed on the page and ishighlighted when that code is active. The DCS
provides access to the different code groups bypositioning the switch to INC or DEC. Sequencing toa different code group does not affect the active
squawks.
AIFF INTERROGATOR
The interrogator provides selective interrogation of
IFF systems along a lineofsight (LOS) or within a
specific scan area. For a specific scan area, theinterrogator may be selectively coupled or decoupledwith the FCR. Refer to T.O. GR1F16CJ3411 fordetails.
Location and classification of a particular IFF system
as a friend or unknown is possible by comparison SIF
codes. During interrogation, the interrogator iscommanded to the modes 1, 2, 3/A, 4 (A or B) code,
PX II and C as programmed via the DTE or manuallyentered using the scratchpad. The mode 4 (A/B)
interrogator code is independent of the mode 4 (A or
B) transponder code as displayed on the CNIinterrogator (INTG) page.
Control of the interrogator is accomplished via the
PX II AUX COMM, PX III IFF control panel, PX II theMISC panel, the PX II UHF VHF IFF transmit, PX III
target management switch, the INTG page, and the
MFDS FCR display. Interrogation of other IFF
systems is possible when in any one of PX II five,
PX III four modes:
D Mode 1 - Security identity.
D Mode 2 - Personal identity.
D Mode 3/A - Traffic identity.
D Mode 4 - Encrypted identity.
D PX II Mode C - Altitude reporting.

1203

T.O. GR1F16CJ1
Operation of the Interrogator on Upfront
Controls PX II
Selecting the INTG page displays and commands
the interrogator to the same mode 1, 2, 3/A, and 4(A/B) codes as the transponder mode 1, 2, 3/A, and 4(A/B) codes. The interrogator continues to use thesame codes until modified via the scratchpad. The
scratchpad is displayed only when data is keyed in andprovides on/off control for each of the five IFF modes
and enterable code control for modes 1, 2, 3/A, and 4.
After modification, the interrogator operates independently of the transponder mode codes. The interrogator is commanded back to the current transponder
mode 1, 2, 3/A, and 4 (A/B) codes via the scratchpad or
by entering the ICP/IKP 7 key via the scratchpad. The
INTG page displays its data on four lines (lines one,
three, four, and five).
The first line displays INTG and the interrogator
power status (ON/OFF) The next area displays thecoupled (CPL) or decoupled (DCPL) status with the
FCR. The area on the right displays the current
selected steerpoint. For CPL and DCPL detailed
description, refer to T.O. GR1F16CJ3411.
The third line displays M1 and enterable mode 1
interrogation code, MC and 5, and the scratchpad. The
5 represents the ICP/IKP key that provides mode Cinterrogation when entered by the scratchpad.
The fourth line displays M2 and enterable mode 2interrogation code, M4 and selectable mode 4 A/B codeand 6. The 6 represents the ICP/IKP key that provides
selection of mode 4 A/B interrogation when entered by
the scratchpad.

AIFF CONTROLS AND INDICATORSIFF IDENT Button
Refer to figure 183. The IFF IDENT button, located on
the instrument panel, provides the primary method of
initiating the identification of position (I/P) function of
the IFF system. Pushing the button momentarily
causes the I/P timer to energize for 1530 seconds. If a
PX II mode 1, 2, 3/A, or S, PX III mode 3/Ainterrogation is recognized within this 1530 second
period, I/P replies are made.
IFF (Mode 4) Caution LightThe IFF (mode 4) caution light, located on the caution
light panel, indicates that the IFF system wasinterrogated on mode 4 and is unable to respondbecause the mode 4 codes (A and B) are zeroized or notcoded, mode 4 is not enabled, or the RF switch is in
QUIET or SILENT.
UHF VHF IFF Transmit Switch PX IIThe UHF VHF IFF transmit switch, located on thethrottle, provides control of the area for interrogation(LOS or scan). For a detailed description, refer to T.O.
GR1F16CJ3411.

IFF IDENT Button(Typical)C

The fifth line displays M3, mode 3/A interrogation
code, RST, and 7. The 7 represents the ICP/IKP key
that resets the interrogator modes and codes to matchthe active transponder modes and codes when enteredby the scratchpad.

DR

Operation of the Interrogator on Upfront
Controls PX III
Selecting either the SCAN INTG page or the LOSINTG page displays the corresponding interrogatormode 1, 2, 3/A, and 4 (A/B) codes. After a DTU load for
IFF codes, the SCAN and LOS interrogator mode 1,
2, 3/A, and 4 (A/B) codes are the same as thetransponder mode 1, 2, 3/A, and 4 (A/B) codes. Modesand codes selected for SCAN are independent of themodes and codes selected for LOS. The interrogator
continues to use the same codes until modified via thescratchpad. For a detailed description, refer to T.O.
GR1F16CJ3411.
1204

DF

1

PX II

1

PX III

1. IFF IDENT Button

Figure 183.

1GR1F-16CJ-1-0082A37

T.O. GR1F16CJ1
Target Management Switch PX III
The target management switch, located on the stick,provides control of the area for interrogation (LOS orscan). For a detailed description, refer to T.O.
GR1F16CJ3411.

INSTRUMENT LANDING SYSTEM (ILS)The ILS provides precision approaches to runways
equipped with localizer, glide slope, and marker
beacon equipment. Localizer identification signals
are supplied to the headset for station identification.
The glide slope and localizer receivers supply glide
slope and localizer deviation data to the deviationbars on the ADI and HUD; the HSI also displays
course or localizer deviation data. Two warning flags,
designated LOC and GS, appear on the ADI when
deviation data is invalid. A course deviation warning
flag appears on the HSI if localizer deviation data isinvalid. HUD symbology consists of localizer andglide slope deviation bars. Dashed deviation bars
indicate invalid data. Deviation bars are rollstabilized with tic marks positioned at the onedot
and twodot deflections.
The symbology automatically displayed on the HUD
with ILS selected is the same as LG down with thefollowing exceptions:
D The great circle steering symbol is replaced by the

ILS deviation bars.

D The lower HUD windows, except distancetodes

tination, are not blanked unless the NLG is lowered
and inertial velocity exceeds 80 knots.

D The

HUD altitude scale does not change from100foot increments to 20foot increments until
NLG is lowered.

D AOA bracket is not displayed until NLG is lowered.

The flight director displays command steering dataon the HUD when selected on the upfront controls.
Command steering symbology consists of a circle, a tic
mark positioned at the top of the circle, and areference mark/caret positioned at the heading/ground track scale. The flight director circle isreferenced to the FPM and appears when localizer
data is valid. Proper use of the flight director requiresthat the localizer be intercepted from a heading no
more than 45 degrees from the localizer course usingbank angles of 30 degrees or less. When the aircraftis within two dots deflection of the localizer deviationbar, the flight director commands a turn to roll out on
the localizer course. The tic mark appears on theflight director circle when glide slope deviation nears
center, indicating that pitch steering data is valid.
The glide slope should be intercepted from a position
that is wings level and on the localizer course. If thepitch steering becomes invalid, the symbol X appears
over the tic mark. The reference caret indicates theheading required to maintain the course selected on
the DED (magnetic heading scale displayed) or
ground track error relative to the course selected(magnetic ground track scale displayed). The course
value may be changed only by entering the new value
through the DED.
The ILS flight director is designed to intercept theglide slope from below while in approximately level
flight. If the aircraft approaches the glide slope fromabove, there is no pitch steering data. The flight
director symbol remains on the horizon, displaying
bank steering data, and the symbol X appears over
the tic mark, indicating that pitch steering data isinvalid. Valid pitch steering is provided after the
glidepath is intercepted.
The marker beacon receiver operates on a fixed
frequency of 75 MHz. Refer to MRK BCN LIGHT, thissection.
ILS Controls
The ILS power knob, located on the AUDIO 2 panel,
controls ILS power and audio volume. The ILS
presentation on the HSI and ADI is controlled by the
CRS set knob (HSI) and INSTR MODE knob. Allother control functions are selected by the upfront
controls. There are no backup controls. In the eventof certain upfront control failures, the ILS may
remain on the last frequency selected prior to failure.

1205

T.O. GR1F16CJ1
Operation of the ILS on Upfront Controls
The ILS is controlled by the upfront controls whenthe ILS power knob on the AUDIO 2 control panel isplaced to on (cw). The ILS page is selected byverifying the CNI page is displayed on the DED andthen depressing the TILS function button. Com
manded (OFF or ON) ILS status and threechangeable items appear:the ILS frequency, thelocalizer course for the ILS flight director cues on the
HUD, and the selection of the flight director (CMD
STRG). The asterisks initialize over the scratchpad
when the TCN/ILS page is selected. The ILS
frequency is changed by insuring that the asterisks
are over the scratchpad, keying in the desired
frequency, and then depressing the ENTR button.
The asterisks then automatically step to the ILS
course for course entry. The flight director isselected/deselected by positioning the asterisks
about the CMD STRG label and depressing theMSEL button. (The flight director is automaticallyselected at ILS power up.) Positioning the INSTR
MODE knob to one of the ILS positions is required
before ILS deviation data (localizer and glide slope)can be displayed on the HSI, HUD, and ADI.
Three steps are required before the ILS flight director(command steering) HUD cues are usable: thelocalizer course is displayed on the DED, the flight
director is modeselected on the upfront controls, andan ILS mode is selected with the INSTR MODE knob.
The course setting on the HSI is not connected to theflight director in the HUD. For consistent ILS
displays, the inbound localizer course should be set on
both the DED and the HSI.

MRK BCN Light
Refer to figure 13. The MRK BCN light is located on
the instrument panel. When the aircraft is over amarker beacon facility, the light illuminates green
and blinks according to the code of the marker beacon.

1206

INSTR MODE SELECT PANEL
Refer to figure 184. The INSTR MODE select panel,
located on the instrument panel, provides for theselection of the displays on the HSI and the ADI.
INSTR HDG Knob C

DF

Refer to figure 184. The INSTR HDG knob, locatedon the INSTR MODE select panel, has arrows
pointing in cw and ccw directions. The knob is pushed
and turned to set the INS heading to a known
magnetic heading in the event of an INS failure(indicated by the AUX warning flag on the ADI).
INSTR MODE Knob
Refer to figure 185 for functions and details of the
INSTR MODE knob, located on the INSTR MODE
select panel. Refer to figure 186 for details of the ADI
and HSI in the instrument modes.
Navigation Aids and Display
Refer to figure 185.
Instrument Modes
Refer to figure 186.

INSTR MODE SelectPanel (Typical)
C

DF

DR

1

1

2
1. INSTR MODE Knob
2. INSTR HDG Knob
GR1F-16CJ-1-0083X37

Figure 184.

T.O. GR1F16CJ1Navigation Aids and Display (Typical)

C

DF

AUDIO 2 CONTROL
PANEL

ADI

DR

AUDIO 2 CONTROL
PANEL

DED

HSI

HUD/ICPCDFAUX COMM PANEL
PX II

IMSP
INSTRUMENT
MODE
SELECTEDRANGE
INDICATOR

RANGE TOTACANSTATIONOR DMENAV AIDNAV

ILS/NAVMRK BCNCRANGE
TO EGI/INS
DESTINATIONCOURSE
ARROW &COURSE
SELECTEDADICOURSEDEVIATION
INDICATORTO FROM
INDICATORMANUALLY
SELECTEDLOCALIZERCOURSELOCALIZERDEVIATIONOUT OF
VIEW

MANUALLY
SELECTEDCOURSEDEVIATION
FROM
SELECTEDCOURSEMANUALLY
SELECTEDLOCALIZERCOURSEDF IFF CONTROL PANELPX IIIHSIILS/TCN

TCN

INSTR
MODE
SELECTPANEL(IMSP)IN VIEWOUT OF
VIEW
LOCALIZERDEVIATION

BEARING
POINTERBEARINGTO TACANSTATIONBEARINGTO EGI/INS
DESTINATION

ATTITUDESPHERE

EGI/INS
ROLL
AND
PITCH
ATTITUDE

LOCALIZERDEVIATION
BAR

GLIDE
SLOPE
DEVIATIONBAR

LOCALIZERDEVIATION

GLIDE
SLOPE
DEVIATIONOUT OFVIEW

OUT OF
VIEW

LOCALIZERDEVIATION

GLIDE
SLOPE
DEVIATION

GR1F-16CJ-1-0084A37

Figure 185.

1207

T.O. GR1F16CJ1

Instrument Modes (Typical)
TCN
NOTE:
The TCN mode has no effecton HUD symbology.

INS/EGI PITCH AND
ROLL ATTITUDE

ADI
EGI/INS MAGNETIC HEADINGMANUALLY SELECTEDCOURSERANGE TO TACANSTATION OR DME
NAV AID

BEARING TO TACANSTATIONTACAN COURSE SIGNALUNRELIABLE

INSTR MODE
SELECT PANEL(IMSP)HEADING REFERENCE MARKER

MANUALLY SELECTEDCOURSE
TO-FROM INDICATOR

DEVIATION FROMSELECTED COURSE

HSI

NAV
NOTE:
The NAV mode has no effecton HUD symbology.

EGI/INS PITCH AND
ROLL ATTITUDE

ADI
EGI/INS MAGNETIC HEADINGMANUALLY SELECTEDCOURSERANGE TO INS/EGIDESTINATIONBEARING TO INS/EGIDESTINATION

EGI/INS MAGNETIC HEADINGSIGNAL UNRELIABLE

INSTR MODE
SELECT PANEL(IMSP)MANUALLY SELECTEDCOURSE
TO-FROM INDICATOR(OUT OF VIEW)

DEVIATION FROMSELECTED COURSE

HSI
Figure 186.(Sheet 1)

1208HEADING REFERENCE MARKER

GR1F-16CJ-1-0085A37

T.O. GR1F16CJ1

Instrument Modes (Typical) PX IIILS/TCN

INS MAGNETIC HEADINGPITCH SCALE (CLIMB)HORIZON LINE

10
CALIBRATEDAIRSPEED20150

AOA BRACKET(NLG LOWERED)

5

SELECTED COURSE( CROSSWIND COMPONENT)(LIMITED)

10
06 065 07

5

02,2

C

10

LOCALIZERDEVIATION

5

5

10

10

02,100

GLIDE SLOPEDEVIATION

02,0

BAROMETRIC ALTITUDE

R 01,450
AL 200150 03

PITCH SCALE (DIVE)

LOCALIZERDEVIATIONROLL SCALEGLIDE SLOPEDEVIATIONLOCALIZER SIGNALUNRELIABLE

GLIDE SLOPE
SIGNALUNRELIABLE

ALTITUDE LOW
SETTINGDISTANCE TO STEERPOINT/STEERPOINT NUMBER

HUD

FLIGHTPATH
MARKER (FPM)

RADAR ALTITUDE

COMMAND STEERING(CMD STRG)NOTES:
1. Glide slope and localizer
bars indicate deviation
only, not steering.
2. Command steering isreferenced to the
localizer course
displayed on the DED.

INS PITCH AND
ROLL ATTITUDE

ADIINS MAGNETIC HEADINGHEADING REFERENCE MARKER

RANGE TO TACANSTATION OR DME
NAV AID

MANUALLY SELECTEDCOURSEBEARING TO TACANSTATIONILS LOCALIZER SIGNALUNRELIABLE

INSTR MODE
SELECT PANEL(IMSP)MANUALLY SELECTEDCOURSE
TO-FROM INDICATOR(OUT OF VIEW)LOCALIZERDEVIATION

HSI

GR1F-16CJ-1-0086A37

Figure 186.(Sheet 2)

1209

T.O. GR1F16CJ1

Instrument Modes (Typical) PX III
ILS/TCN

EGI MAGNETIC HEADINGPITCH SCALE (CLIMB)HORIZON LINE

10
CALIBRATEDAIRSPEED20150

AOA BRACKET(NLG LOWERED)LOCALIZERDEVIATION

5

SELECTED COURSE( CROSSWIND COMPONENT)(LIMITED)

10
06 065 07

5

02,2

C

10
0.88 5
4.1
ILS
090 25 10OWNSHIP BEARING
AND RANGE FROMBULL'S-EYE

ALTITUDE LOW
SETTING

AL 2005

02,100

GLIDE SLOPEDEVIATION

02,0

BAROMETRIC ALTITUDE

R 01,450
F008.005:50
035 05

10

DISTANCE TO STEERPOINT/STEERPOINT NUMBER

HUD

FLIGHTPATH
MARKER (FPM)

PITCH SCALE (DIVE)

LOCALIZERDEVIATIONROLL SCALEGLIDE SLOPEDEVIATIONLOCALIZER SIGNALUNRELIABLE

GLIDE SLOPE
SIGNALUNRELIABLE

RADAR ALTITUDE

COMMAND STEERING(CMD STRG)NOTES:
1. Glide slope and localizer
bars indicate deviation
only, not steering.
2. Command steering isreferenced to the
localizer course
displayed on the DED.

EGI PITCH AND
ROLL ATTITUDEADIEGI MAGNETIC HEADINGHEADING REFERENCE MARKER

RANGE TO TACANSTATION OR DME
NAV AID

MANUALLY SELECTEDCOURSEBEARING TO TACANSTATIONILS LOCALIZER SIGNALUNRELIABLE

INSTR MODE
SELECT PANEL(IMSP)MANUALLY SELECTEDCOURSE
TO-FROM INDICATOR(OUT OF VIEW)LOCALIZERDEVIATION

HSI
Figure 186.(Sheet 3)1210

GR1F-16CJ-1-1086X37

T.O. GR1F16CJ1

Instrument Modes (Typical) PX IIILS/NAV

INS MAGNETIC HEADINGPITCH SCALE (CLIMB)HORIZON LINE

10
CALIBRATEDAIRSPEED20150

AOA BRACKET(NLG LOWERED)

5

SELECTED COURSE( CROSSWIND COMPONENT)(LIMITED)

10
06 065 07

5

02,2

C

10

LOCALIZERDEVIATION

5

5

10

10

02,100

GLIDE SLOPEDEVIATION

02,0

BAROMETRIC ALTITUDE

R 01,450
AL 200150 03

PITCH SCALE (DIVE)

LOCALIZERDEVIATIONROLL SCALEGLIDE SLOPEDEVIATIONLOCALIZER SIGNALUNRELIABLE

GLIDE SLOPE
SIGNALUNRELIABLE

ALTITUDE LOW
SETTINGDISTANCE TO STEERPOINT/STEERPOINT NUMBER

HUD

FLIGHTPATH
MARKER (FPM)

RADAR ALTITUDE

COMMAND STEERING(CMD STRG)NOTES:
1. Glide slope and localizer
bars indicate deviation
only, not steering.
2. Command steering isreferenced to the
localizer course
displayed on the DED.

INS PITCH AND
ROLL ATTITUDE

ADIINS MAGNETIC HEADINGHEADING REFERENCE MARKER
MANUALLY SELECTEDCOURSERANGE TO INSDESTINATIONBEARING TO INSDESTINATIONILS LOCALIZER SIGNALUNRELIABLE

INSTR MODE
SELECT PANEL(IMSP)MANUALLY SELECTEDCOURSE
TO-FROM INDICATOR(OUT OF VIEW)LOCALIZERDEVIATION

HSI

GR1F-16CJ-1-0087A37

Figure 186.(Sheet 4)

1211

T.O. GR1F16CJ1

Instrument Modes (Typical) PX III
ILS/NAV

EGI MAGNETIC HEADINGPITCH SCALE (CLIMB)HORIZON LINE

10
CALIBRATEDAIRSPEED20150

AOA BRACKET(NLG LOWERED)LOCALIZERDEVIATION

5

SELECTED COURSE( CROSSWIND COMPONENT)(LIMITED)

10
06 065 07

5

02,2

C

10
0.88 5
4.1
ILS
090 25 10OWNSHIP BEARING
AND RANGE FROMBULL'S-EYE

ALTITUDE LOW
SETTING

AL 2005

02,100

GLIDE SLOPEDEVIATION

02,0

BAROMETRIC ALTITUDE

R 01,450
F008.005:50
035 05

10

DISTANCE TO STEERPOINT/STEERPOINT NUMBER

HUD

FLIGHTPATH
MARKER (FPM)

PITCH SCALE (DIVE)

LOCALIZERDEVIATIONROLL SCALEGLIDE SLOPEDEVIATIONLOCALIZER SIGNALUNRELIABLE

GLIDE SLOPE
SIGNALUNRELIABLE

RADAR ALTITUDE

COMMAND STEERING(CMD STRG)NOTES:
1. Glide slope and localizer
bars indicate deviation
only, not steering.
2. Command steering isreferenced to the
localizer course
displayed on the DED.

EGI PITCH AND
ROLL ATTITUDEADIEGI MAGNETIC HEADINGHEADING REFERENCE MARKER
MANUALLY SELECTEDCOURSERANGE TO EGIDESTINATIONBEARING TO EGIDESTINATIONILS LOCALIZER SIGNALUNRELIABLE

INSTR MODE
SELECT PANEL(IMSP)MANUALLY SELECTEDCOURSE
TO-FROM INDICATOR(OUT OF VIEW)LOCALIZERDEVIATION

HSI
Figure 186.(Sheet 5)

1212

GR1F-16CJ-1-1087X37

T.O. GR1F16CJ1

Instrument Modes (Typical)
ILS HUD DISPLAYSINBOUND LOCALIZER COURSE = 090

RUNWAY
7

6

45

8

9

10

5
2
01

015

02

45 INTERCEPT HEADING(090 - 45 = 045
AIRCRAFT HEADING)
4
CMD STRG SYMBOL APPEARS (NO PITCHSTEERING PROVIDED)LOCALIZER DEVIATION BAR IS SOLID (VALID)A RIGHT TURN IS COMMANDED (CMD STRGSYMBOL IS TO RIGHT OF FPM)

3
2

3
02

023

03

30

BANK ACHIEVED

1
INITIAL AIRCRAFTHEADING = 015

1

4
01

015

02

04

NO CMD STRG SYMBOLLOCALIZER AND GLIDE SLOPE DEVIATIONBARS ARE DASHED (INVALID)COURSE CARET IS LIMITEDHORIZON LINE IS DISPLAYED FOR
REFERENCE

045

AIRCRAFT IS ON 45

05

INTERCEPT HEADING

GR1F-16CJ-1-0088X37

Figure 186.(Sheet 6)

1213

T.O. GR1F16CJ1

Instrument Modes (Typical)
ILS HUD DISPLAYS

5

8
04

045

08 08305

A RIGHT TURN IS COMMANDED (CMD STRG SYMBOL IS TO
RIGHT OF FPM)LOCALIZER DEVIATION BAR INDICATES APPROXIMATELY 2
DOTS LEFT
6

09

AFTER TIC MARK APPEARS ON TOP OF CMD STRG SYMBOL,PITCH STEERING IS PROVIDED
AIRCRAFT IS ON LOCALIZER COURSE AND ABOVE GLIDESLOPE
906 06208 08307

30 BANK ACHIEVED
LOCALIZER DEVIATION BAR INDICATES APPROXIMATELY
1 DOT LEFT
709

AIRCRAFT IS ON LOCALIZER COURSE AND GLIDE SLOPE(ALL SYMBOLS ALIGNED)
10

08 0830907

AIRCRAFT IS ON LOCALIZER COURSE (CMD STRG SYMBOL,COURSE CARET, AND LOCALIZER DEVIATION BAR ARE ALL
ALIGNED)COURSE CARET INDICATES 083 VS 090 HEADING DUE TO
CROSSWIND

080

09

AIRCRAFT IS LEFT OF LOCALIZER COURSE AND GLIDE SLOPEIS INVALID (CMD STRG SYMBOL HAS X ON IT. NO PITCHSTEERING IS PROVIDED)

GR1F-16CJ-1-0089X37

Figure 186.(Sheet 7)1214

T.O. GR1F16CJ1FLIGHT INSTRUMENTS
Refer to figure 187. The flight instruments arelocated on the instrument panel. The instruments
listed below are common and are not illustrated indetail.
D Airspeed/Mach Indicator.
D Clock.
D Magnetic Compass.
D ServoPneumatic Altimeter.
D Standby Attitude Indicator.
D Vertical Velocity Indicator.

ALTIMETER
The servopneumatic altimeter is a dual modepressure altitude indicator with a range of -1000 to+80,000 feet. The operating mode is manually
selected by the mode lever located at the lower rightcorner of the instrument. In the ELECT (primary)operating mode, the altimeter is electrically driven by
the CADC. In the PNEU (secondary) operating mode,
the altimeter is pneumatically operated by staticpressure supplied by the pitotstatic system. Should
the CADC or altimeter servo malfunction, thealtimeter automatically reverts to the pneumatic
mode and the PNEU flag appears on the face of thealtimeter. The PNEU flag may also appear when
accelerating or decelerating through the transonicregion or while performing high g maneuvers.
The barometric setting knob, located at the lower left
corner of the instrument, is used to set the desired
altimeter setting. C DF The barometric setting of thealtimeter is electrically transmitted to the CADC asa manual input correction for the pressure altitudedisplay on the HUD. The barometric setting is shownin inches of mercury (Hg).

Flight Instruments (Typical)
1

C

DF

DR

1

2
2

10
9
8

3
4

10
9
8

5

3
4

6

7
7

1. AOA Indexer
2. Standby Attitude Indicator
3. Altimeter

6
4. Vertical Velocity Indicator
5. Magnetic Compass
6. Clock

7. Horizontal Situation
Indicator
8. Attitude Director Indicator

9. AOA Indicator
10. Airspeed/Mach Indicator
GR1F-16CJ-1-0090X37

Figure 187.

1215

T.O. GR1F16CJ1
AIRSPEED/MACH INDICATORVERTICAL VELOCITY INDICATOR (VVI)

The airspeed/mach indicator is pneumaticallyoperated by total and static pressure supplied by thepitotstatic system. The indicator displays indicatedairspeed and mach number. Indicated airspeed isdisplayed by a moving airspeed/mach pointer against
a fixed airspeed scale. Mach number, which is readagainst the airspeed/mach pointer, is displayed by arotating mach scale. The range of the indicator is from80850 knots and from 0.52.2 mach.

The VVI displays rate of climb/descent information
provided by the CADC. The indicator has avertically moving tape display with a range of 6000
fpm climb or dive. The VVI is powered by emergency
ac bus No. 1.
MAGNETIC COMPASS

The maximum allowable airspeed pointer indicates
800 knots at sea level. Higher airspeeds are indicatedas altitude is increased. This indication is not a valid
maximum allowable airspeed cue; it should bedisregarded. Refer to MAXIMUM AIRSPEEDOPERATING LIMITATIONS, Section V.

The magnetic compass is a selfcontained indicatorwhich shows the heading of the aircraft with respect
to magnetic north. Adjustable compensating magnets
in the compass provide the means for cancellingmagnetic disturbances originating within the aircraft. A deviation correction card for the compass islocated immediately below and aft of the compass.

The SET INDEX knob is used to set the airspeed
reference index.

ATTITUDE DIRECTOR INDICATOR (ADI)

STANDBY ATTITUDE INDICATOR (SAI)The SAI is a selfcontained, electrically poweredvertical gyroscope that mechanically positions theattitude sphere of the indicator to display aircraft
pitch and roll attitudes. Manual caging of thegyroscope is accomplished by pulling the PULL TO
CAGE knob at the lower right corner of the indicator.
The knob is held out until the sphere is caged to zero
pitch and roll indication and then released. Adjust
ment of the miniature aircraft reference symbol isaccomplished by rotation of the PULL TO CAGE knob.
Since the SAI is mounted in the instrument panel atan angle, it indicates a pitch angle of 4 degrees lessthan the ADI when pitch trim knobs on bothindicators are set at the pitch trim index. If caging isrequired, the aircraft should be flown wings level,
constant altitude, and at an AOA of approximately+4 degrees. When caged on the ground, allow 2
minutes prior to taxi.
An OFF warning flag appears whenever electrical
power is lost or whenever the PULL TO CAGE knob
is pulled. After power loss, the indicator continues toprovide usable attitude information for approxi
mately 9 minutes. The gyroscope of the indicator isunrestricted in roll but is limited to approximately"85 degrees in pitch.
The indicator can develop errors during aerobatic
maneuvering, primarily when pitch is near 90
degrees. If these errors exceed 7 degrees after
returning to level flight, erection is cut off. If thisoccurs, the gyro does not automatically erect andmust be manually caged to eliminate the error. The
indicator is powered by battery bus No. 2.
1216

Refer to figure 188. The ADI displays pitch and roll
attitude information supplied by the EGI/INS. The
ADI is not limited in pitch or roll and displays any
aircraft attitude accurately. In certain modes ofoperation, the indicator displays ILS glide slope andlocalizer deviation information. The instrument
displays turn rate which is presented in standard turnneedle format. The turn rate needle is driven by therate gyroscope transmitter which senses the aircraft
turn rate and displaces one needle width in response
to a 11/2degree/second turn rate. The slip indicator(ball) is a selfcontained item. The pitch trim knob isused to adjust the attitude sphere to the desired pitch
attitude in reference to the miniature aircraft.
The OFF warning flag may indicate failure of either
the EGI/INS or the ADI. A momentary OFF and/or
AUX flag, even with proper attitude being displayed,
may indicate impending failure of the ADI or
EGI/INS data to the ADI. The GS warning flag
indicates that the glide slope deviation bar isunreliable. The LOC warning flag indicates that the
localizer signal is unreliable. The AUX warning flag
signifies that the EGI/INS has failed or is operating
in a less precise attitude condition and that HSI
heading must be set to a known heading by the C DF
INSTR HDG knob on the INSTR MODE select panel.
HORIZONTAL SITUATION INDICATOR (HSI)
Refer to figure 189. The HSI displays horizontal or planview of the aircraft with respect to the navigation
situation. The miniature aircraft symbol in the center ofthe HSI is fixed and comparable to an aircraft
superimposed on a compass rose. The face of the HSI isa compass card driven by the EGI/INS so that aircraft
magnetic heading is always read at the upper lubber line.

T.O. GR1F16CJ1The HDG set knob provides the means for rotating theheading reference marker to the desired heading. Once
set, the heading reference marker rotates with thecompass card. The heading reference marker provides
a reference to the heading select mode of the autopilot.
The CRS set knob provides the means for selecting
any one of 360 courses. To select a desired course,
rotate the head of the course arrow to the desired
course on the compass card and check the course
selector window for the precise setting. Once set, thecourse arrow rotates with the compass card.
PX III The EGI does not use slaved variation forcourse display for TACAN, VORTAC, and VOR points
retrieved from the database. Rather, the magnetic
variation at the aircraft present position is used forcourse computation and display.
PX III Except for UTM's and markpoints (steerpoints
2130), each steerpoint can have an associated CDI
deflection. the selected deflection scales the HSI for
course deviation as represented by a twodot
deflection of the CDI. From the STPT L/L or DEST
DIR pages, DCS SEQ can be used to display a DED
page to program these functions. Two options for CDI
lateral deflection are available: 4 NMI (default) or 3
DEG. The 3degree selection will be displayed in thelower left corner of the HUD as ANG. UTM's andmarkpoints always use 4 nm scaling.

The bearing pointer provides bearing information to
TACAN station or EGI/INS destination. Refer to
NAVIGATION AIDS AND DISPLAY, this section.
The range indicator provides a readout of distance in
nm to a TACAN station, DME navigational aid, or
EGI/INS destination. Loss of TACAN or DME signal
or an unreliable signal causes a warning flag to cover
the range indication window when ILS/TCN or TCN
is selected. When NAV or ILS/NAV is selected, an
EGI/INS failure causes the warning flag to cover therange indication window. Loss of power to the HSI
causes the OFF warning flag to come into view.

containing stored flight data: the Signal Acquisition
Unit (SAU) and the Crash Survivable Memory Unit
(CSMU). The mishap investigation data, or Type 1
data, is stored in the CSMU. Service life data (Types
2 and 3 data), engine usage data (Type 4 data), andavionics health and diagnostic data (Type 5 data) arestored in the SAU. The SAU is not crash hardened.
Service life and engine usage data is stored forapproximately the last seven flights. Type 5 data isavailable for the last two or three flights. Mishap
investigation data storage is limited to the lastapproximately 60 minutes of flight time; however,there are automatic special event provisions in flightthat protect an approximate 30 seconds worth of datafrom overwrite until the next flight. If desired, amanual special event data save can be accomplishedby depressing the FIRE & OHEAT DETECT test
button for 1 second. Saved data covers at least the 15
seconds prior to and after button depression. A total often special events (automatic and manual) may besaved in Type 1 data. There is no limit to the numberof special events that can be stored in Type 5 data. The
occurrence of a special event also generates an MFL.
Special events are:D Acceleration (NZ) greater than 9.5g or less than

-3.5g.

D AOA greater than 29 degrees or less than or equal

to -5 degrees.

D Canopy unlocked in flight.
D Engine data save.
D Engine overspeed.
D Engine overtemperature.
D Engine stagnation.
D Engine subidle.
D Engine transfer to SEC.

CLOCK

D EPU overspeed.

The clock, located on the right auxiliary console, is an
8day, manually wound clock with provisions for anelapsed time indication up to 60 minutes.

D FLCS reset.

CRASH SURVIVABLE FLIGHT DATA RE
CORDER (CSFDR)
The aircraft contains a CSFDR system that storesengine usage data, aircraft service life data, avionics
health and diagnostic data, and mishap investigation
data. The CSFDR system consists of two units

D FLCS warning light.
D LEF's locked by asymmetry brakes.
D Main fuel shutoff valve not fully open.
D Main generator off line.
D OVERHEAT caution light.
D Throttle in OFF in flight.

1217

T.O. GR1F16CJ1

Attitude Director Indicator (ADI) (Typical)
1

2

3

4

5

19

6
7

18

8

17

9
16
15

14

13 12

11

10

Glide Slope Unreliable
1. GS Warning Flag
2. Attitude Sphere
3. Pitch Scale
4. Upper Bank Index Pointer5. Localizer Deviation Bar
Localizer Signal Unreliable
6. LOC Warning Flag
INS/EGI Navigation and7. AUX Warning Flag
Heading Unavailable
8. Pitch Trim Index
9. Pitch Trim Knob
10. Rate-of-Turn Scale
11. Rate-of-Turn Needle
12. Slip Indicator (Ball)
13. Lower Bank Index Pointer14. Lower Bank Scale
15. Miniature Aircraft Symbol
16. Glide Slope Deviation Bar and PointerAttitude Sphere/INS/
17. OFF Warning Flag
EGI Unreliable
18. Glide Slope Deviation Scale
19. Horizon Line
1F-16X-1-0013X

Figure 188.

Horizontal Situation Indicator (HSI)(Typical)
1

2

3

4

5

19

6

18

7

17

8

16

9

15

10

14

13

12

1. Range Indicator
2. Warning Flag
Range Indicator
3. Heading Reference Marker
4. Upper Lubber Line
5. Course Selector Window
6. Bearing Pointer7. OFF Warning Flag
HSI Power
8. TO-FROM Indicator
9. Miniature Aircraft Symbol
10. CRS Set Knob
11. Course Arrow
Tail
12. Bearing PointerTail
13. Lower Lubber Line
14. HDG Set Knob
15. Course Deviation Indicator
16. Course Deviation Scale
17. Warning Flag
Course Deviation
18. Course Arrow
19. Compass Card11
1F-16X-1-0014X

Figure 189.

1218

T.O. GR1F16CJ1

F16D AIRCRAFT
Only those items which are different, significant, or
unique to the forward or rear cockpit are discussed in
the following paragraphs.
ENGINE CONTROLS DR
Only a throttle, ENG CONT switch, and FUELMASTER switch are provided.
Throttle DR
The throttle is limited in certain functions:
D Cannot be advanced from OFF to IDLE.
D Cannot be advanced from MIL to AB.
D Cannot be retarded from IDLE to OFF.
D Does not have a throttle friction control.

ENG CONT Switch DR
Refer to figure 13. The ENG CONT switch, located on
the left console, is a twoposition switch.
Functions are:
D SEC - Manually selects SEC (SEC caution lightilluminates) regardless of forward cockpit ENG
CONT switch position.

D NORM - (Guarded position) Returns control toforward cockpit switch position.

FUEL SYSTEM
Fuel Control Panel DR
Refer to figure 123. The fuel control panel, located on
the left console, contains a guarded FUEL MASTER
switch. The FUEL MASTER switches in both cockpits
must be in the MASTER (on) position to permit fuelflow to the engine. Either switch when positioned to
OFF shuts off all fuel flow to the engine.
SPEEDBRAKE SYSTEM
SPD BRK Switch DR
Refer to figure 116. The SPD BRK switch isspringloaded to the off (center) position and must be
held in either the open (aft) position or close (forward)
position.

DRAG CHUTE SYSTEMDRAG CHUTE Switch DR
Refer to figure 141. The DRAG CHUTE switch,
located on the LG control panel, is a threeposition
guarded switch used to deploy and release the dragchute. Manually holding the DRAG CHUTE switch in
REL prevents drag chute deployment when the DF
switch is placed to DEPLOY.
Manually holding the DRAG CHUTE switch in REL
after DF drag chute deployment (switch remaining in
DEPLOY) allows the deployed drag chute to release
if airspeed is above 60 knots.
Manually holding the DRAG CHUTE switch in
DEPLOY prevents drag chute release when the DF
switch is moved from DEPLOY to NORM/REL.
LANDING GEAR (LG) SYSTEMDR All LG functions are duplicated except that the
BRAKES channel, ANTISKID, and LANDING TAXILIGHTS switches are not present. In addition, the
DN LOCK REL button function is different.

DN LOCK REL Button DR
Refer to figure 137. The DN LOCK REL button,
located on the LG control panel, when depressed,
unlocks the LG handle electrically to allow movement
to UP in case the left MLG WOW switch has failed to
the ground position. The DN LOCK REL button has
no mechanical unlock function.
FLIGHT CONTROL SYSTEM
Simultaneous inputs to the forward and aft sticks (orrudder pedals) are added together to position theflight control surfaces accordingly. Unintentionalright roll inputs can occur as a result of leg/knee
interference with the stick.
DR The flight controls are abbreviated and contain
only the stick, rudder pedals, and warning, caution,and status lights. The FLCP and MANUAL TRIM
panels are not available.

STICK CONTROL OPERATIONWith the STICK CONTROL switch in FWD, the stick
indicator displays the word FWD. Depressing theforward cockpit paddle switch locks out the rear
cockpit stick, rudder, and MPO commands. The
OVRD light in both cockpits illuminates indicating
override is activated. Depressing the rear cockpit
paddle switch has no effect on the forward cockpit
controls.
1219

T.O. GR1F16CJ1
With the STICK CONTROL switch in AFT, the stick
indicator displays the word AFT. Depressing therear cockpit paddle switch, PX III with the ASIU
switch OFF or the ASIU STICK CONT switch in
FLIGHT, locks out the forward cockpit stick, rudder,and MPO commands and transfers control of NWS tothe rear cockpit. The OVRD light in both cockpits
illuminates, indicating override is activated. De
pressing the forward cockpit paddle switch has noeffect on the rear cockpit controls.

Functions are:
D NORM
S Activation of ejection system from rear cockpit

results in canopy jettison, then a .33second delay
followed by only rear seat being ejected.
Activation of ejection system from forward cockpit
following rear seat ejection results in forward seatejection.

S Activation of ejection system from forward cockpit

results in canopy jettison, a .33second delay, aftseat ejection, and a .4second delay followed byforward seat ejection.

STICK CONTROL Switch DF
Refer to figure 118. The STICK CONTROL switch
is located on the TEST switch panel.
Functions are:

D AFT - Activation of ejection system from either

cockpit results in canopy jettison, a .33second
delay, aft seat ejection, and a .4second delay
followed by forward seat ejection.

D SOLO - Activation of ejection system from forward
D FWD - The aft cockpit flight control functions arelocked out as long as the paddle switch is held
depressed.

D AFT - The forward cockpit flight control functions

are locked out, and NWS control is transferred to
the rear cockpit as long as the paddle switch is held
depressed.

Stick Indicator DR
Refer to figure 190. A threeposition stick indicator,
located on the instrument panel, indicates the
position of the STICK CONTROL switch bydisplaying the word AFT or FWD as applicable.
Diagonal stripes appear when power is removed orappear momentarily during switching.

cockpit results in canopy jettison, then a .33second
delay followed by only forward seat being ejected.
Activation of ejection system from aft cockpit
results in canopy jettison, then a .33second delay
followed by only aft seat being ejected. Simulta
neous activation from forward and aft cockpits
results in unsequenced ejections.

Stick Indicator and
OVRD Light D(Typical)PX III

Stick OVRD Lights D
Refer to figure 190. The stick OVRD lights, locatedon the forward and rear cockpit instrument panels,
illuminate when the paddle switch is used to takecontrol.

DF

1

PX II

2
DR

ESCAPE SYSTEM DEJECTION MODE SEL Handle DR

1. OVRD Light (Amber)
2. Stick Indicator
GR1F-16CJ-1-0093A37

Refer to figure 155. The EJECTION MODE SEL
handle is located on the right auxiliary console.
1220

Figure 190.

T.O. GR1F16CJ1COMMUNICATIONS SYSTEM
Refer to figure 179. Identical audio control panels,
with fully active switches, are located in both
cockpits. The UHF, VHF, TACAN, and ILS can be
enabled from either cockpit with the audio control
panels but must be turned off in both cockpits todeactivate the equipment. Squelch functions only if
squelch is enabled in both cockpits. Control of otherfunctions is determined by the position of the DF
PX II CNI, PX III C & I knob.
Functions are:
D UFC - Switches on the upfront controls are fully
active in DF and DR . Control inputs can be madefrom either or both cockpits; however, only one
variable (e.g., UHF frequency) can be selected orchanged at one time. Only one DED, presented inboth cockpits, can be selected at a time.

D BACKUP - DR

No control is available for mode,
code, channel, or frequency of the UHF, TACAN, or
IFF. Audio control panels are not affected.

INSTR MODE SELECT PANEL D
Refer to figure 184. If the INSTR MODE knob is in
the same position in both cockpits, the rear cockpit
HSI course slaves to that selected in the front cockpit.
DR The INSTR MODE select panel does not have an
INSTR HDG knob.
FLIGHT INSTRUMENTS DR
An accelerometer is installed and the altimeterbarometric pressure set knob is not connected to the
CADC.

stick and handson control inputs from analog signals
to digital signals for transmission to the SMS andselected stores. Refer to figure 179.1 for handsoncontrols located on the throttle and stick.
The switches on the ASIU panel have the following
functions:
STICK CONTROL Switch DR PX III
The STICK CONTROL switch is a twoposition
guarded switch. The switch positions are marked
FLIGHT and AVIONICS.
Functions are:
D FLIGHT - Allows normal flight functions for aft

stick and handson controls when the ASIU switch
is in ASIU and the DF STICK CONTROL switch isin FWD.

D AVIONICS - Allows expanded avionic functions foraft stick and handson controls when the ASIU
switch is in ASIU and the DF STICK CONTROL
switch is in FWD. Allows normal flight operations foraft stick and handson controls when ASIU switch isin OFF.

Accelerometer DR(Typical)1
4

Accelerometer DR
Refer to figure 191. The accelerometer is selfcon
tained and mechanically indicates acceleration
acting along the vertical axis of the aircraft. The
accelerometer is graduated in g units and has three
indicating pointers. The main pointer displays
instantaneous changes in acceleration. The positive
and negative auxiliary pointers indicate the maxi
mum positive and negative acceleration experienced.
The auxiliary pointers retain their highest readings
until the PUSH TO SET knob is depressed.
AFT STATION INTERFACE UNIT (ASIU) DR PX III
Refer to figure 192. The ASIU allows the aft stick to
perform either regular flight control functions or
expanded avionic functions. The ASIU panel islocated on the right console, to the right of the aftstick. When activated, the ASIU converts aft station

3

2

1. Auxiliary Positive Pointer
2. Auxiliary Negative Pointer3. PUSH TO SET Knob
4. Main PointerGR1F-16CJ-1-0094X37

Figure 191.

1221

T.O. GR1F16CJ1

ASIU Panel DR PX III
1

ARMT CONSENT position. The switch is in series
with the MASTER ARM switch in the forward cockpit
and must be in the ARMT CONSENT position to
enable the normal release of any store.
AFT SEAT HUD MONITOR (ASHM) DR

2

Refer to figure 195. The HUD CTVS video can bedisplayed on the ASHM or on the right MFD,
depending on the position of the VIDEO SEL switch.
SERVICING DIAGRAM
Refer to figure 196 for servicing/specifications
information.

1. STICK CONTROL Switch
2. ASIU Switch

GR1F-16CJ-1-1097X37

Throttle and Stick
Switches/Controls
DR PX II

Figure 192.
SWITCH/CONTROLASIU Switch DR PX III
The ASIU switch is a twoposition lever lock switch.

STATUS

THROTTLE
DOG FIGHT/Missile Override

Inactive/Inactive

Functions are:

SPD BRK

Active*RDR CURSOR/ENABLE

Active/Active

D ASIU - Power on.

*ANT ELEV

ActiveD OFF - Power off.

*MAN RNG/UNCAGE

Active/Active

UHF/VHF/IFF

Active/Active/Inactive

For a detailed system description, refer to T.O.
GR1F16CJ3411.
MULTIFUNCTION DISPLAY (MFD) DR
Refer to figure 13. The HUD CTVS video can bedisplayed on the right MFD by placing the VIDEO
SEL switch to HUD.
THROTTLE AND STICK SWITCHES/CONTROLS DR
Refer to figures 193 and 194 for the switches/con
trols located on the throttle and stick.
ARMT CONSENT SWITCH DR
Refer to figure 13. The ARMT CONSENT switch,
located on the instrument panel, is guarded in the1222

STICK
TRIM

ActiveCAMERA/GUN

Active/Inactive

WPN REL

Active*Target Management

Active*Display Management

ActiveNWS A/R DISC

Active/Active

MSL STEP

ActivePaddleActiveExpand/FOV

ActiveCountermeasures Management

Inactive* Control gained by first upward activation of display management
switch.

Figure 193.

T.O. GR1F16CJ1

Throttle and Stick Switches/Controls
DR PX IIISTATUSSWITCH/CONTROLTHROTTLE
DOG FIGHT/Missile Override
SPD BRKRDR CURSOR/ENABLE
ANT ELEVMAN RNG/UNCAGE
Transmit (UHF/VHF/IFF)
STICK
TRIM
CAMERA/GUN
WPN REL
Target Management
Display Management
NWS A/R DISCMSL STEP
PaddleExpand/FOVCountermeasures Management

ASIU POWER ON

ASIU POWER OFF

Inactive/Inactive
ActiveActive/Active
ActiveActive/Active
Active/Active/Active/Inac
tive

Inactive/Inactive
ActiveInactive/Inactive
ActiveInactive/Inactive
Active/Active/Active/Inac
tive

ActiveActive/Inactive
ActiveActiveActiveActive/Active
ActiveActiveActiveInactive

ActiveActive/Inactive
ActiveInactive
Inactive
Active/Active
Inactive
ActiveInactive
Active (aft CMSonly)/Inactive

FWD

NOTES:
D Switch/control functions vary depending on selection of FLIGHT or AVIONICS on the ASIU panel.
D Refer to T.O. GR1F16CJ3411.

Figure 194.

Change 11223

T.O. GR1F16CJ1

Aft Seat HUD Monitor DR
2

1

1. VIDEO SEL Switch
2. HUD Monitor
3. Handgrip
4. BRT Control
5. Function Knob
6. CONT Control

3

4

5

6
GR1F-16CJ-1-0095X37

Figure 195.

1224

T.O. GR1F16CJ1

Servicing Diagram (Typical)
COCKPIT (EJECTION SEAT(S))
EMERGENCY GASEOUS
OXYGEN BOTTLE(S)NLG WHEEL WELL
LG/HOOK EMERGENCY
PNEUMATIC BOTTLE(NITROGEN)

DOOR 2318
EXTERNAL POWER CABLE
RECEPTACLE

DOOR C 2105, D 2107
GROUND COOLING
RECEPTACLE

DOOR 3306
IFF MODE 4PANEL 3408 (TOP)EPU (MONOPROPELLANT)DOOR 2317
ENGINE AND EMS GONO-GO INDICATORSDOOR 3206
AMMUNITION LOADINGPANEL 3402 (TOP)LEF POWER DRIVE UNIT

DOOR 3308
PX II LIQUID OXYGENDOOR 3115
HYDRAULIC SYSTEM B
GROUND TEST ANDSERVICE CONNECTIONSPANEL 3412 (TOP)EPU OIL SERVICE

DOOR 3101
HYDRAULIC SYSTEM B
ACCUMULATOR

DOOR 3216
HYDRAULIC SYSTEM AGROUND TEST ANDSERVICE CONNECTIONSLMLG WHEEL WELL
HALON RESERVOIRDOOR 3202
HYDRAULIC SYSTEM A
ACCUMULATOR

DOOR 3103
GROUND REFUELING
RECEPTACLE

DOOR 3208
EPU (NITROGEN)HYDRAZINE LEAK
DETECTORDOOR 3204
DEFUELING
RECEPTACLE
PW229 / GE129
DOOR 3218/3320
ENGINE OIL SIGHT GAGE

DOOR 3318
ADG
CSD
BOTTOM VIEW
DOOR 4113
FLCS ACCUMULATORS (2)DOOR 3316
ENGINE OIL SERVICE
JOAP SAMPLE

DOOR 4304
BRAKE/JFSACCUMULATORS (2)DOOR 4214
DRAG CHUTE ACCUMULATOR

GR1F-16CJ-1-0096A37

Figure 196.(Sheet 1)

1225

T.O. GR1F16CJ1

Servicing Diagram (Typical)
SPECIFICATIONSSERVICEABLE ITEM

FUEL

OIL

HYDRAULIC FLUIDOXYGENUSAF

ENGINE/JFS

NATO

MILT5624, JP4

F40

MILT5624, JP5

F43 OR F44

MILT83133, JP8F34

JET A, B (COMMERCIAL)NONEJET A1 (COMMERCIAL)

F35

PW229 ENGINE *MILL7808J OR LATER

GE129 ENGINE *MILL7808

ADG/CSD/EPU

MILL7808

0148

HYDRAULIC SYSTEMS A AND B

MILH5606
MILH83282

H515
H537

GASEOUSMILO27210, TYPE I

PX IIMILO27210, TYPE II

LIQUID

0 148
0148

NONE

EXTERNAL ELECTRICAL
POWER

115 (15) VAC, 400 (30) HZ

A/M32A60ANONE

NITROGENGASEOUSBBN441A, TYPE I, GRADE BNONEFUEL TANK INERTING
AGENT (OPTIONAL)

LIQUID

HALON 1301

NONE

MONOPROPELLANT (EPU)

LIQUIDHYDRAZINE (70% N2H4, 30%
H2O)

NONE

* IF NECESSARY, ENGINE LUBRICATING OILS MILL7808 (NATO CODE 0148) AND MILL23699 (NATO CODE 0156) MAY BE MIXED. AT
THE FIRST OPPORTUNITY THEREAFTER, THE OIL SHALL BE DRAINED AND FLUSHED, AND THE ENGINE SERVICED WITH THE PROPER
LUBRICATING OIL AS SPECIFIED IN THE APPLICABLE ENGINE TECHNICAL ORDERS.

Figure 196.(Sheet 2)1226

T.O. GR1F16CJ1

SECTION II
NORMAL PROCEDURESTABLE OF CONTENTS
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Flight Planning . . . . . . . . . . . . . . . . . . . . . . . .
Takeoff and Landing Data Card . . . . . . . . .
Cockpit Designation Code . . . . . . . . . . . . . . .
Weight and Balance . . . . . . . . . . . . . . . . . . . .
Preflight Check . . . . . . . . . . . . . . . . . . . . . . . . . .
Exterior Inspection . . . . . . . . . . . . . . . . . . . . .
Cockpit Access . . . . . . . . . . . . . . . . . . . . . . . . .
Before Entering Cockpit . . . . . . . . . . . . . . . .
Cockpit Interior Check . . . . . . . . . . . . . . . . .
Left Console . . . . . . . . . . . . . . . . . . . . . . . . .
Left Auxiliary Console . . . . . . . . . . . . . . .
Instrument Panel . . . . . . . . . . . . . . . . . . . .
Right Auxiliary Console . . . . . . . . . . . . . .
Right Console . . . . . . . . . . . . . . . . . . . . . . .
After Cockpit Check Is
Complete  Verify . . . . . . . . . . . . . . . . . . . .
Before Starting Engine . . . . . . . . . . . . . . . . . . . .
Starting Engine PW 229 . . . . . . . . . . . . . . . . . . .
Starting Engine GE 129 . . . . . . . . . . . . . . . . . . .
After Engine Start . . . . . . . . . . . . . . . . . . . . . . .
Before Taxi . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Taxi . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Before Takeoff . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Takeoff . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Normal Takeoff . . . . . . . . . . . . . . . . . . . . . . . .
Takeoff With Asymmetric Stores . . . . . . . . .
Climb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Climb/InFlight/Operational Checks . . . . . . . .
Air Refueling Procedures . . . . . . . . . . . . . . . . . .
Descent/Before Landing . . . . . . . . . . . . . . . . . . .
Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Normal Landing . . . . . . . . . . . . . . . . . . . . . . .
Short Field Landing (Dry Runway) . . . . . .
Crosswind Landing . . . . . . . . . . . . . . . . . . . . .
TouchandGo Landing . . . . . . . . . . . . . . . . .
After Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Prior to Engine Shutdown . . . . . . . . . . . . . . . . .
Engine Shutdown . . . . . . . . . . . . . . . . . . . . . . . .
Scramble . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Preflight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Aircraft On Quick Response Status . . . . . .
Hot Refueling . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Hot Refueling Precautions . . . . . . . . . . . . . .
Prior to Hot Pit Entry . . . . . . . . . . . . . . . . . .

Page
21
21
21
21
21
22
22
22
22
23
24
25
25
25
26
26
27
28
210
211
217
218
218
220
220
222
222
222
223
223
223
223
226
226
227
227
228
229
230
230
230
231
231
231

Prior to Hot Refueling . . . . . . . . . . . . . . . . . .
During Hot Refueling . . . . . . . . . . . . . . . . . . .
Hot Refueling Complete . . . . . . . . . . . . . . . .
Quick Turnaround . . . . . . . . . . . . . . . . . . . . . . . .
Prior to Engine Shutdown . . . . . . . . . . . . . .
Special Procedures . . . . . . . . . . . . . . . . . . . . . . .
EPU Hydrazine Support at NonF16
Bases . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Supplemental Procedures . . . . . . . . . . . . . . . . .
ILS Procedures . . . . . . . . . . . . . . . . . . . . . . . .
Exterior Inspection . . . . . . . . . . . . . . . . . . . . .
Danger Areas . . . . . . . . . . . . . . . . . . . . . . . . . .
Turning Radius and Ground Clearance . . . .
GSuit Hose Routing . . . . . . . . . . . . . . . . . . .
Strange Field Procedures . . . . . . . . . . . . . . . . . .

231
232
232
232
232
232
232
232
232
232
232
232
232
232

INTRODUCTION
This section provides the actions required for normal
operation of the aircraft. Amplification is included
only when special considerations or techniquesshould be observed. A complete knowledge of Section
III, EMERGENCY PROCEDURES, and Section V,
OPERATING LIMITATIONS, is required prior toflight.
FLIGHT PLANNING
Refer to T.O. GR1F16CJ11.
TAKEOFF AND LANDING DATA CARD
A takeoff and landing data card is provided in the
Flight Crew Checklist for recording computed data.
COCKPIT DESIGNATION CODE
An asterisk (*) preceding steps is used to highlightprocedures for D aircraft which apply to both
cockpits.
WEIGHT AND BALANCE
For maximum GW limitations, refer to Section V,
OPERATING LIMITATIONS. For weight andbalance information, refer to the individual aircraft
Form F (DD Form 3654) and the Weight and Balance
Handbook.
21

T.O. GR1F16CJ1

PREFLIGHT CHECKCheck AFTO Form 781 for aircraft release and storesstatus.
EXTERIOR INSPECTION
Refer to figure 24 for normal preflight inspection.

Cockpit Access(Typical)
EXTERNAL CANOPY SWITCH(LOWER SURFACE OF LEFTSTRAKE)

COCKPIT ACCESS
Refer to figure 21.
1. Canopy-Open by positioning external cano
py switch to the up position.

If winds exceed 30 knots, open the
canopy only as far as needed to enter/exit
the cockpit. Decreasing the canopy angle
reduces the possibility that the canopy
can be blown past full open.

1F-16X-1-1045A

Figure 21.

2. Ladder-Position on cockpit sill.

S Survival KIT DEPLOYMENT switch-As
desired.

C A failure of the canopy actuator
could allow the canopy to fall duringtransit. Do not position the ladderuntil the canopy motion has stopped.

NOTE
Cockpit access is gained by the useof two entrance ladders or moving one
ladder as required.
D

BEFORE ENTERING COCKPIT*1. Ejection seat-Check.

S Ejection safety lever-Safe (up).
S Safety pins (2)-Removed (ejection safety

lever and EMERGENCY MANUAL CHUTE
handle).

22

A Parachute Landing Fall (PLF) withthe survival kit not deployed may resultin injury.

S RADIO BEACON-As desired.
S Emergency oxygen bottle-1800 psi mini
mum.

S Quickdisconnect (left side)-Connected.
S Recovery parachute-Free from damage andgrease.

S Inertia reel straps retaining pin (yellow)-Visible in circular inspection hole.

S Parachute risers and SEAWARS -Free fromdamage and grease.

T.O. GR1F16CJ1
8. Audio panels-Set:

S COMM 1 mode knob-SQL.
Rapidly pulling on the shoulder
harness to check the locking mecha
nism may cause damage to the locking
mechanism.

S All other knobs-CCW.

S Lanyards from canopy to seat-Check.

9. ALT GEAR handle-In.

S Emergency

10. ALT FLAPS switch-NORM.

oxygennect-Connected.

hose

quickdiscon

11. GND JETT ENABLE switch-OFF.

S Environmental sensor pitot tubes-Clear of

obstructions and if flipup pitot tubes are
installed, safety pins (2) removed.

12. DRAG CHUTE Switch-NORM.
13. HOOK switch-UP.

S Quickdisconnect (right side)-Connected.

14. ARMT CONSENT switch-ARMT CONSENT (guard down).

S Electronic recovery sequencer battery indi

15. EJECTION MODE SEL handle-SOLO.

cator-White indication.

S CANOPYJETTISONsafety pin removed.

Thandle-Secure,

S Lapbelt retaining pins (yellow)-Protrud
ing, both sides.

S Lapbelts-Secure. (Pull up and forward,
both sides.)S Survival kit straps-Secure. (Pull up andaft, both sides.)2. MAIN PWR switch-OFF.
DR

S COMM 2 mode knob-SQL.

For solo flight:

3. Ejection seat-Safe, straps secure, pins
removed.
4. CANOPY JETTISON
safety pin removed.

Thandle-Secure,

5. SPD BRK switch-Center.
6. FUEL MASTER switch-MASTER (guarddown).
7. ENG CONT switch-NORM (guard down).

16. Interior LIGHTING
knobs off.

control

panel-All

17. OXYGEN REGULATOR-OFF and 100%.
18. Utility light-OFF and secured.
COCKPIT INTERIOR CHECKThe checklist steps are not arranged in a mandatoryorder.
*1. Loose or foreign objects-Check.
A thorough cockpit interior preflight checkshall be accomplished prior to each flightwith emphasis on loose or foreign objects
that might cause injury to personnel ordamage to the aircraft.
*2. Harness and personal equipment-Fasten.
Attach the parachute risers to the harness.
Attach and adjust survival kit straps.
Secure and adjust the lapbelt. After
fastening lapbelt, hold right side of lapbelt
buckle stationary and shake and push left
side. Also insure that latching mechanism
has reset back to its original position and isflush with buckle to insure a positive lock.
Connect oxygen hose, gsuit hose, vest hose,and communication leads. Check operationof the shoulder harness locking mechanism.

23

T.O. GR1F16CJ1
3. FLCS PWR TEST switch-NORM.
4. DEFOG lever-Midrange.
FA partially locked lapbelt may open

during maneuvering flight or ejection
sequence.

FUnobserved gsuit hose or vest hose

disconnects can contribute to gray/blackout or loss of consciousness.

FFailure to adjust survival kit straps toachieve a snug fit between the pilot
and seat pan lid may result in injury
during ejection.

FFailure to adequately secure andtighten lapbelt may result in inability
to reach and operate the MPO switch
during outofcontrol situations.

5. DIGITAL BACKUP switch-OFF.
*6. ALT FLAPS switch-NORM.
7. MANUAL TF FLY UP switch - ENABLE.
8. LE FLAPS switch-AUTO.
9. BIT switch-OFF.
10. TRIM/AP DISC switch-NORM.
11. ROLL, YAW, and PITCH TRIM-Center.
*12. FUEL MASTER switch-MASTER (guard
down and C DF safetywired).
13. TANK INERTING switch-OFF.
14. ENG FEED knob-NORM.

Do not adjust seat height with survival
kit straps or lapbelt disconnected asdamage to the ejection seat or stick
may occur.
NOTE
FRefer to figure 211. The recommended

routing of the gsuit hose is directly
under the torso harness and aft of thesurvival kit strap to reduce thepossibility of a gsuit hose disconnect.
Gsuit hose routing must provide
sufficient slack to allow for maximum
mobility since inflight reconnect isextremely difficult.

FExcess gsuit hose must be properly

routed to prevent MANUAL TRIM
panel interference during flight.

F DR

Incorrectly routed or unsecuredgsuit hose may result in throttleinterference or disconnected gsuit hose.

*3. Rudder pedals-Adjust.
Adjust rudder pedals so that legs are flat on
the seat cushion to prevent the right leg
from hitting stick or injury during ejection.
Left Console
1. PROBE HEAT switch-OFF.
2.

STICK CONTROL switch-As briefed
when DR occupied; FWD for solo flight.
DF24

15. AIR REFUEL switch-CLOSE.
16. IFF MASTER knob-STBY.
17.

PX II CNI, PX III C & I knob-BACKUP.

*18. TACAN-As desired.
19. EXT LIGHTING control panel-As required.

PX III If ground operation of the
anticollision strobe light is anticipatedto be longer than 2 hours, the
ANTICOLLISION knob should be set
in position 1 to avoid overheating thepower supply and strobe.

20. MASTER light switch-NORM.
21. EPU switch-NORM (guard down).
22. MAIN PWR switch-OFF.
23. AVTR power switch-OFF.
24. AVTR PX II DISPLAY, PX III VIDEO SELECT
knob-HUD.
25. ECM power-Off.
*26. COMM 1 power knob-CW.
*27. COMM 1 mode knob-SQL.

T.O. GR1F16CJ1*28. COMM 2 power knob-CW.

4.

*29. COMM 2 mode knob-SQL.
*30. TACAN power knob-CW.

DR ARMT CONSENT switch-ARMT CONSENT (guard down).

5. LASER ARM switch-OFF.
*6. DRAG CHUTE switch-NORM.

31.

C DF AB RESET switch-NORM.

32.

C DF ENG CONT switch-PRI (guard down).

*7. HUD/ASHM-Set.

33.

DR ENG CONT switch-NORM (guard down).

*8. Altimeter-Set.

34. JFS switch-OFF.

9. FUEL QTY SEL knob-NORM.

35. UHF radio backup control panel:

10. EXT FUEL TRANS switch-NORM.

a. Function knob-BOTH.
b. Frequency-As desired.
36. Throttle-Verify freedom of motion, then OFF.

NOTE
FIf a wing transfer valve fails and WING

FIRST is selected, fuel flows onto theground from the overboard vent line
under the left wing (outboard of thefuel pylon). Selecting NORM shouldstop the fuel spillage.

*37. SPD BRK switch-Forward.
*38. DOG FIGHT switch-Center.
Left Auxiliary Console

F PX III If CFT's are installed and an

external tank transfer valve fails toclose when the tank is empty and
WING FIRST or NORM is selected,
fuel is trapped in the CFT and sometrapped fuel may vent overboard.
Selecting CFT FIRST/NO FILL should
prevent trapped/vented CFT fuel.

*1. ALT GEAR handle-In.
2. CMDS switches (9)-OFF.
3. RF switch-NORM.
4.

PX III SYMBOLOGY power knob-Off.

5. STORES CONFIG switch-As required.
6. LANDING TAXI LIGHTS switch-OFF.
*7. LG handle-DN.
*8. GND JETT ENABLE switch-OFF.
9. BRAKES channel switch-CHAN 1.

*11. INSTR MODE knob-As desired.
Right Auxiliary Console*1. Clock-Set.
2.

DR EJECTION MODE SEL handle-b2p
NORM or AFT (as briefed), LESS b2p AFT.

10. ANTISKID switch-ANTISKID.
*11. EMER STORES JETTISON button-Coverintact.
*12. HOOK switch-UP.
Instrument Panel
1. ROLL switch-ATT HOLD.
2. PITCH switch-A/P OFF.
3. MASTER ARM switch-OFF.

FSOLO position should not be selectedwith the rear seat occupied.

F DR Ejection initiated with the EJECTION MODE SEL handle in NORM
may incapacitate the DF pilot atairspeeds greater than 180 knots. Useof the NORM position is therefore notrecommended.

Change 125

T.O. GR1F16CJ1*14.

Right Console

S Pressure-Check 50120 psi.

NOTE

S Mode lever-PBG/ON (as required).

Oxygen regulator pressure de
creases slowly when OBOGS is notoperating and may indicate low prior
to engine start.
PX IIIS Diluter lever-NORM.
S EMER lever-NORM.

1. SNSR PWR switches (4)-OFF.

S FLOW indicator-Check.

2. HUD control panel-Set.

S EMER lever-EMER.

3. NUCLEAR CONSENT switch-OFF (guarddown).

Check for positive oxygen pressure andmask and hose/connector leakage.

4. ZEROIZE switch-OFF.

S EMER lever-NORM.

*5. Wristrest and armrest-As desired.
*6. Interior
desired.
7.

LIGHTING

PX II Oxygen system-Check:control

panel-As

C DF VOICE MESSAGE switch-VOICEMESSAGE.
If safety wire is broken, notify maintenance.

AFTER COCKPIT CHECK IS COMPLETE-VERIFY
The following items are those important switches
that, if not correctly positioned, could cause a safety
hazard and/or improperly operated systems during
engine start.

8. TEMP knob-AUTO.

*1. FUEL MASTER switch-MASTER (guard
down and C DF safetywired).

9. AIR SOURCE knob-NORM.

2. ENG FEED knob-NORM.
3. EPU switch-NORM (guard down).

If AIR SOURCE knob is in OFF or
RAM, electronic equipment may bedamaged.

4.

C DF ENG CONT switch-PRI (guard down).

5.

DR ENG CONT switch-NORM (guard down).

*6. Throttle-OFF.
*7. LG handle-DN.

10. AVIONICS POWER switches (8)-OFF.
11. ANTI ICE switch-AUTO/ON.

If engine antiice is on or has been on,heat in the inlet strut may be sufficient
to cause injury if touched.

Insure that the LG handle is confirmed
fully down. The LG handle can be in anintermediate position allowing LG
extension and/or safe indications;however, the LG handle is not locked
and LG retraction could occur during
subsequent ground operations.
*8. HOOK switch-UP.

If there is visible moisture andambient temperature is 45_F (7_C) or
less, place the ANTI ICE switch to ON.
12. IFF ANT SEL switch-NORM.
13. UHF ANT SEL switch-NORM.
26

D If either HOOK switch is in DN, thehook extends when aircraft battery
power is applied.

9. MASTER ARM switch-OFF.

T.O. GR1F16CJ1
10. AIR SOURCE knob-NORM.
*11. Loose or foreign objects-Check.

Loose or foreign objects in the cockpit
can cause entanglement or obstruction
with critical controls (e.g., ejection
handle, stick, rudder pedals, throttle,
canopy handle, etc.).

F PX III Labored breathing through theoxygen mask before engine start ispossible with the diluter lever in
NORM. It is not possible to to breath
through the oxygen mask before
engine start with the diluter lever in
100%.

2. FLCS PWR TEST switch-TEST and hold.
Activation of the FLCS PWR TEST switch
with the MAIN PWR switch in BATT closes
FLCS relays and allows verification of
power output of the FLCC with the aircraft
battery as the power source.

S Verify lights on:
BEFORE STARTING ENGINE
1. MAIN PWR switch-BATT.

S ACFT BATT TO FLCS.
S FLCS PMG.
S FLCS PWR (4).

S Verify FLCS RLY light off.
To prevent possible depletion of bat
tery power, do not allow MAIN PWR
switch to remain in BATT or MAIN
PWR for more than 5 minutes without
engine running or external power
applied.
NOTEFA JFS RUN light flashing once persecond indicates a noncritical failurein the JFS system. If the JFS operates
normally for engine start, the mission
may be continued. Notify maintenanceafter flight.

FA JFS RUN light flashing twice persecond indicates a critical failure in
the JFS system. The JFS will not
operate. Place the MAIN PWR switch
to OFF and notify maintenance.

S Verify FLCS RLY light on.
NOTE
FToe brakes are not available when the

FLCS RLY light is on and the aircraft
battery is the only operating powersource.

3. FLCS PWR TEST switch-Release.
4. MAIN PWR switch-MAIN PWR.

S Verify lights on:
S ELEC SYS.
S HYD/OIL PRESS.
S FLCS RLY.
S SEC.
S ENGINE.

Do not apply external electrical power
for more than 30 minutes without
cooling air; longer operation may
damage electronic components whichcannot be turned off. Prior to applying
external power without cooling air,insure that all avionic equipment isoff; UHF radio operation on backup
control is permissible.
27

T.O. GR1F16CJ1
NOTE
If FLCS RLY light is off, cycle MAINPWR switch to BATT and back to
MAIN PWR to open FLCS relays.
5. EPU GEN and EPU PMG lights-Confirm
off.

If either light is illuminated, turn the
MAIN PWR switch to OFF. Insure that
the EPU safety pin remains installed
and notify maintenance. (The EPU
will activate using hydrazine if the
EPU safety pin is removed.)
6. Communications-Established.
7. Canopy-As desired.

C A failure of the canopy actuator
could allow the canopy to fall duringtransit. Keep hands and arms out ofthe path of canopy travel duringopening or closing.

FTo prevent possible engine damage,stow or secure loose cockpit items prior
to engine start with the canopy not
closed.

F C DF When lowering or raising the

canopy handle, insure that the canopy
switch is not bumped to the up (open)
position before handle is completely
locked or unlocked. If this occurs, holdthe switch in the down (close) position
as soon as possible to relieve thejammed condition and prevent canopy
actuator motor burnout. If required,manually crank canopy toward theclosed position to relieve the jammed
condition. If possible, place MAINPWR switch to OFF and have theaircraft battery disconnected by theground crew. This action removes
power from the canopy actuator.

28

8. Chocks in place, fireguard posted, and intake
and other danger areas clear (ground crew).

STARTING ENGINE PW 229
Refer to figure 28 for danger areas. Moving the JFS
switch to either start position closes the FLCS
relays. The FLCS RLY light goes off; the FLCS PMG
light (and ACFT BATT TO FLCS light if engine isstarted on battery power) illuminates. Toe brakes
also become functional (if engine is started onbattery power).
The JFS RUN light comes on within 30 seconds; after
the JFS RUN light illuminates, engine rpm starts toincrease smoothly. At approximately 15 percent
rpm, the hydraulic and fuel pressures increase. The
SEC caution light goes off by approximately 18
percent rpm. The throttle should be advanced to
IDLE after engine rpm reaches 20 percent.
Immediately after advancing the throttle to IDLE,
the SEC caution light illuminates for about 3
seconds to indicate automatic DEEC/SEC groundselftest. Engine lightoff occurs within 20 seconds
after throttle advance and is indicated by airframe
vibration and an increase in rpm followed by anincrease in FTIT. Without external power connected,
only the RPM and FTIT indicators function until thestandby generator is on line.
At approximately 50 percent rpm, the JFS automati
cally shuts down. At approximately 55 percent rpm,
the standby generator comes on line, the ENGINE
warning light goes off, and the OXY LOW warning
light illuminates for approximately one second. To
insure the emergency buses are being powered by thestandby generator, check for illumination of the
SEAT NOT ARMED caution light and three green
WHEELS down lights prior to the main generator
coming on line. Five to ten seconds after the thestandby generator comes on line, the main generator
comes on line, and the standby generator goes off line.
FTIT increases smoothly at a moderate rate andnormally peaks at less than 625_C.
1. JFS switch-START 2.
2. SEC caution light-Check off.

T.O. GR1F16CJ1
3. Throttle-Advance to IDLE at 20 percent
rpm minimum.

If the FUEL MASTER switch is
discovered in OFF during start (i.e.,
after the throttle is moved to IDLE),
discontinue the start. Positioning the
FUEL MASTER switch to MASTER
with the throttle out of OFF may cause
a hot start or tailpipe fire.
NOTE
FIf the HYD/OIL PRESS warning lightgoes off and stays off prior toapproximately 25 percent rpm, suspect failure of the engine oil pressure
sensor and abort the aircraft. This
check is not valid if less than 30
seconds elapse from movement of theMAIN PWR switch out of OFF until
both hydraulic pressures are above
1000 psi (i.e., low oil pressure timedelay has not expired; in which case,the HYD/OIL PRESS warning lightgoes off when hydraulic pressuresexceed 1000 psi, but may reilluminate
if the 30second time delay expires
while oil pressure is still low.

FIf the main and/or standby generator

fails to come on, do not attempt reset.
Although the generator(s) may reset, aproblem may still exist. Notify
maintenance.

FThe SEC caution light illuminates

FIf an engine restart is desired shortly

after engine shutdown, motor theengine with the JFS for a minimum of
1 minute or a maximum of 4 minutes
before throttle advancement to insure
the FTIT drops below 200_C. Initiat
ing a start with the FTIT above 200_C
may result in a DEEC FTIT limited
hung start.

4. ENGINE warning light-Offmately 55 percent rpm).
*Engine at idle and check:
5. JFS switch-Confirm OFF.

If JFS switch does not automatically
return to OFF, turn JFS off. Notify
maintenance after flight.
6. HYD/OIL PRESS warning light-Off.
The HYD/OIL PRESS warning light may
not go off until rpm is increased to 70
percent. If the warning light comes on again
at idle, notify maintenance.
7. FUEL FLOW - 5001500 pph.
8. OIL pressure-15 psi (minimum).
9. NOZ POS-Greater than 80 percent.

immediately after throttle advance forabout 3 seconds indicating automatic
DEEC/SEC ground selftest.

10. RPM-6577 percent.

FIf the FLCS RLY light remains on with

11. FTIT-625_C or less.

the JFS switch in either start position,
one or more FLCS relays did not latch
closed. Latching may occur when thestandby or main generator comes online. If the FLCS RLY light goes off
when either generator comes on line,
notify maintenance of the occurrence
after flight. If the FLCS RLY lightremains on after the main generator ison line, notify maintenance.

(approxi

12. HYD PRESS A & B-28503250 psi.
13. Six fuel pump lights (ground crew)-On.
NOTE
FFP light may blink at idle rpm.

29

T.O. GR1F16CJ1
14. Main fuel shutoff valve (ground crew)- Check.

Do not make stick inputs while groundcrew is in proximity of control surface.
15. JFS doors (ground crew)-Verify closed.
16. Throttle cutoff release-Check.
Verify that the cutoff release does notremain in the actuated position by attempt
ing to retard the throttle to OFF without
depressing the cutoff release. If throttle
moves to OFF, remain in OFF and notify
maintenance.
NOTE
Failure to perform this check afterengine start can result in an unde
tected stuck throttle cutoff release,which may lead to an unintentional
engine shutdown.

STARTING ENGINEGE 129Refer to figure 28 for danger areas. Moving the JFS
switch to either start position closes the FLCS
relays. The FLCS RLY light goes off; the FLCS PMG
light (and ACFT BATT TO FLCS light if engine isstarted on battery power) illuminates. Toe brakes
also become functional (if engine is started onbattery power).
The JFS RUN light comes on within 30 seconds; after
the JFS RUN light illuminates, engine rpm starts toincrease smoothly. The throttle should be advanced
to IDLE after engine rpm reaches either 20 percent
rpm (ambient temperature less than 90_F) or
maximum motoring rpm (ambient temperature
90_F or greater). Engine lightoff occurs within 10
seconds after throttle advance and is indicated by anairframe vibration and an increase in rpm followedby an increase in FTIT. Without external power
connected, only the RPM and FTIT indicators
function until the standby generator is on line.
The SEC caution light goes off at 20 percent rpm. At
approximately 55 percent rpm, the JFS automati
cally shuts down. At approximately 60 percent rpm,
the standby generator comes on line and the
ENGINE warning light goes off. To insure theemergency buses are being powered by the standby
generator, check for illumination of the SEAT NOTARMED caution light and three green WHEELS
210

down lights prior to the main generator coming online. Five to ten seconds after the standby generator
comes on line, the main generator comes on line andthe standby generator goes off line.
FTIT increases smoothly at a rapid rate.
1. JFS switch-START 2.
NOTE
Abnormal rpm indications (i.e. instantaneous jumps) may indicate alterna
tor failure and not rpm indicatormalfunction. Abort the aircraft andnotify maintenance.
2. Throttle-Advance to IDLE at either 20
percent rpm minimum (ambient temperature less than 90_F) or maximum motoringrpm (ambient temperature 90_F or greater).

If the FUEL MASTER switch is
discovered in OFF during engine start(i.e., after the throttle is moved toIDLE), discontinue the start. Position
ing the FUEL MASTER switch to
MASTER with the throttle out of OFF
may cause a hot start or tailpipe fire.
NOTE
FIf the HYD/OIL PRESS warning lightgoes off and stays off prior toapproximately 15 percent rpm, suspect failure of the engine oil pressure
sensor and abort the aircraft. This
check is not valid if less than 30
seconds elapse from movement of theMAIN PWR switch out of OFF until
both hydraulic pressures are above
1000 psi (i.e., low oil pressure timedelay has not expired; in which case,the HYD/OIL PRESS warning lightgoes off when hydraulic pressuresexceed 1000 psi, but may reilluminate
if the 30second time delay expires
while oil pressure is still low).

FIf the main and/or standby generator

fails to come on, do not attempt reset.
Although the generator(s) may reset, aproblem may still exist. Notify mainte
nance.

T.O. GR1F16CJ1
NOTE
If the FLCS RLY light remains on with
the JFS switch in either start position,
one or more FLCS relays did not latch
closed. Latching may occur when thestandby or main generator comes online. If the FLCS RLY light goes off
when either generator comes on line,
notify maintenance of the occurrence
after flight. If the FLCS RLY lightremains on after the main generator ison line, notify maintenance.
3. SEC caution light-Off.
4. ENGINE warning light-Off (approximately
60 percent rpm).

14. Main fuel shutoff valve (ground crew)-Check.

Do not make stick inputs while groundcrew is in proximity of control surface.
15. JFS doors (ground crew)-Verify closed.
16. Throttle cutoff release-Check.
Verify that the cutoff release does notremain in the actuated position by attempt
ing to retard the throttle to OFF without
depressing the cutoff release. If throttle
moves to OFF, remain in OFF and notify
maintenance.
NOTE

NOTE
In hot day conditions (temperature
greater than 90_F (32_C)) the EN
GINE warning light may not go off
until idle speed is obtained (greaterthan 70 percent rpm).
*Engine at idle and check:
5. JFS switch-Confirm OFF.

If JFS switch does not automatically
return to OFF, turn JFS off. Notify
maintenance after flight.
6. HYD/OIL PRESS warning light-Off.
The HYD/OIL PRESS warning light may
not go off until rpm is increased to 70
percent. If the warning light comes on again
at idle, notify maintenance.
7. FUEL FLOW- 7001700 pph.
8. OIL pressure-15 psi (minimum).
9. NOZ POS-Greater than 94 percent.
10. RPM-6280 percent.
11. FTIT-650_C or less.
12. HYD PRESS A & B-28503250 psi.
13. Six fuel pump lights (ground crew)-On.
NOTE
FFP light may blink at idle rpm.

FFailure to perform this check afterengine start can result in an unde
tected stuck throttle cutoff release,which may lead to an unintentional
engine shutdown.

FEngine

vibration/rumble may besensed on the ground while accelerating or decelerating the engine, whentransferring to SEC, or when stabi
lized at idle in either PRI or SEC. Thisvibration/rumble has no adverse effecton the engine or aircraft structure. The
vibration/rumble should disappear
when the throttle is advanced slightly(approximately 5 percent rpm increase).

AFTER ENGINE START
1. TEST switch panel-Check:a. PROBE HEAT switch - PROBE HEAT.
PROBE HEAT caution light off. If thecaution light illuminates, one or moreprobe heaters are inoperative or amonitoring system failure has occurred.
b. PROBE HEAT switch - TEST.
PROBE HEAT caution light flashes 35times per second. If the caution light
does not illuminate or if it illuminates
but does not flash, the probe heatmonitoring system is inoperative.
c.

PROBE HEAT switch - OFF.

d. FIRE & OHEAT DETECT button-Test.
211

T.O. GR1F16CJ1e.

PX II OXY QTY test switch-Test andcheck.

f.

MAL & IND LTS button-Test.
NOTE

FProper VMS operation is verified by

3.

PX III AVIONICS POWER panel-Set:

a. MMC switch-MMC.
b. ST STA switch-ST STA.
c.

MFD switch-MFD.

d. UFC switch-UFC.

the presence of each word in priority
sequence (i.e., PULLUP, ALTITUDE,WARNING, etc.). A brief LG warning
horn is also heard prior to the
WARNING and CAUTION words.

4. EGI/INS-Align ( PX III after display visibleon the DED).

F D Only the LG warning horn and theNOTE

warning and caution voice messages
are heard when MAL & IND LTS test
is initiated from the rear cockpit.

FLESSD
b7n PX III
The circuitry
powering the AVIONICS FAULT caution lights and ECM lights is notcapable of turning on both the DF andthe DR lights except during MAL &IND LTS test. As a result, the only timethese lights illuminate in either
cockpit is during MAL & IND LTS test.
For avionics faults, failure of the
AVIONICS FAULT caution lights tocome on also prevents illumination of
the MASTER CAUTION light andoperation of the voice caution message.
Faults will still be displayed on the
PFLD. The interim solution is removalof the light bulbs from the DR
AVIONICS FAULT caution light andDR ECM light. This solution enables
operation of the DF AVIONICS FAULT
caution light, DF ECM light, both
MASTER CAUTION lights, and thevoice caution message after a fault
occurs. The DR AVIONICS FAULT
caution light and DR ECM light should
not illuminate during MAL & IND LTS
test if the bulbs have been removed. If
the DR AVIONICS FAULT caution
light or DR ECM light illuminates
during MAL & IND LTS test, contact
maintenance.

2.

e. GPS TRK switch-GPS TRK.

F PX III If GPS is unavailable, failure toenter alignment coordinates flags thealignment as degraded (NAV RDY
light does not flash).

F PX II Entry of alignment coordinates isrequired even if internal coordinates
are exactly equal to parking spot
location.

FFailure to enter alignment coordinates

flags the alignment as degraded (NAV
RDY light does not flash).

5. SNSR PWR panel:a. LEFT HDPT
required.

switch-OFF,

Select OFF unless TFR/FLIR opera
tion is planned. With LEFT HDPT
power on and the TFR in OFF/STBY, afailure in the WX mode circuitry cancause the TFR to operate in the WX
mode. An unexpected flyup may
subsequently occur.
b. RIGHT HDPT switch-As required.
c.

FCR switch-FCR.

PX II AVIONICS POWER panel-Set:

a. FCC switch-FCC.

d. UFC switch-UFC.

In environments of high humidity,
delay turning the FCR on for approxi
mately 4 minutes. This action pre
cludes damage to FCR caused bymoisture in the ECS immediately afterengine start.

e. GPS switch-GPS.

d. RDR ALT switch-RDR ALT.

b. SMS switch-SMS.
c.

MFD switch-MFD.

212Change 1

unlessT.O. GR1F16CJ1S ENG CONT switch-SEC.

*6. HUD/ASHM-As desired.
7.

PX II CNI, PX III C & I knob-UFC.

S NOZ POS - Less than 5 percent.

NOTE

S SEC caution light-On.

If UFC battery failure occurs, UHF/VHF preset frequencies in the UFC
revert to default values until enteredmanually or by the DTC.

S Throttle- Verify engine response tothrottle movement, then IDLE.

S ENG CONT switch- C DF PRI, DR NORM.

8. MFL-Clear.

S NOZ POS - Greater than 80 percent.

NOTE
Multiple nonresettable FLCS MFL's
may occur during engine start and may
be accompanied by a lack of control
surface response to stick/rudder commands. These MFL's can occur if
aircraft battery voltage is low or if aswitching problem occurs when the JFS
switch is moved out of OFF. If multiple
FLCS MFL's remain after a FLCS reset:
b1q Perform the following:
D DBU switch-BACKUP, then OFF.

Cycling the DBU switch withWOW reinitializes the FLCC andinhibits FLCS MFL reporting for
20 seconds.

If FLCS MFL's clear and do not recur
after 20 seconds, maintenance action
is not required.
LESS b1q Perform the following:
D DBU switch - BACKUP, then OFF.

10.

GE 129 SEC-Check. May be delayed until
the BEFORE TAKEOFF check.

S ENG CONT switch-SEC.
S SEC caution light-On.
S RPM-Stabilized.

RPM may drop up to 10 percent from PRI
value before stabilizing. Stabilized SEC
idle rpm may be up to 5 percent lower
than that in PRI.

S Throttle-Snap to MIL and then snap to

IDLE when rpm reaches 85 percent. Checkfor normal indications and smooth operation.

S NOZ POS-10 percent or less within 30

seconds after selecting SEC.

S ENG CONT switch- C DF PRI, DR NORM.

D FLCS RESET switch - RESET.

S SEC caution light-Off.

If FLCS MFL's clear, maintenanceaction is not required.

S NOZ POS-Greater than 94 percent.

If FLCS MFL's remain:
D FLCS BIT - Initiate and monitor.

If FLCS MFL's remain:
D Engine - Shut down, then restart tocycle power to the FLCC.

If FLCS MFL's clear following FLCS BITor engine restart, notify maintenance ofthe occurrence after flight.
9.

S SEC caution light-Off.

PW 229 SEC-Check after the engine has
run at idle for at least 30 seconds. May bedelayed until the BEFORE TAKEOFF check.

S Throttle-IDLE.

11. Flight controls-Cycle.
NOTE
To assist in warming the hydraulicfluid and removing air from thehydraulic system, maximum stick andrudder pedal inputs should be madeprior to running FLCS BIT.
12. FLCS BIT-Initiate and monitor.
Position BIT switch to BIT. The RUN lighton FLCP illuminates. At successful comple
tion of BIT (approximately 45 seconds), theRUN light goes off, the BIT switch returns
to OFF, and the FAIL light and FLCS
warning light remain off. A BIT passmessage appears on the FLCS MFD page.
Change 1213

T.O. GR1F16CJ1
NOTE
FIf the FLCS BIT reports a failure

through the FLCS warning light andthe FAIL light on the FLCP, the failure
cannot be reset. The BIT must be
reinitiated. In this case, the RUN lightand the FAIL light are simultaneously
illuminated for the first steps of the BIT,
after which the FAIL light goes off
unless BIT detects a subsequent failure.

FIf the PX II FCC, PX III MMC is notturned on prior to FLCS BIT, an FLCSMUX DEGR message appears on the
PFLD.

FEntering system altitude during FLCS

BIT results in an incorrect system
altitude.

FDuring FLCS BIT, the ECS air ejectors

are turned off. This results in higher
ECS component temperatures andmay cause fumes in the cockpit that
could be described as metallic odors.
This situation is normal and does not
require maintenance action.

F PW 229 During FLCS BIT, the CADC

provides a mach signal to the engine
that results in an rpm increase of
approximately 2 percent.

13. ECM panel-As required.
14. SPD BRK switch-Cycle.

S RSVR-Each reservoir indicates approxi
mately 460/480 pounds.

S PX II INT WING-Each wing indicatesapproximately 525/550 pounds.

S PX III INT WING & CFT - If CFT's are notinstalled (or installed and not filled), eachwing indicates approximately 525/550
pounds. If CFT's are filled, each pointerindicates approximately 2000/2100 pounds.

S EXT WING - Each external wing tankindicates approximately 2300/2420 pounds(370gallon fuel tank) or 3750/3925 pounds(600gallon tank) (for full tanks).

S EXT CTR-FR pointer indicates approxi

mately 1800/1890 pounds (for full tank). AL
pointer drops to zero.
NOTE

The sum of the individual fuel tanks
and the totalizer should agree within"100 pounds with only internal fuel or"300 pounds with external fuel.

S FUEL QTY SEL knob-As desired.
18. EPU FUEL quantity-95102 percent.
19. Avionics-Program as required and verify(manual or data transfer cartridge).

S SMS.

*15. WHEELS down lights-Three green.
*16. SAI-Set.
17. FUEL QTY SEL knob-Check.
The following values are based on JP4 or JP5/8:

S TEST-FR, AL pointers indicate 2000("100) pounds and totalizer indicates 6000("100) pounds. FWD and AFT FUEL LOW
caution lights illuminate. PX III If CFT's are
installed, CFT pumps operate and shut down
30 seconds after switch is moved out of TEST
(if INT WING & CFT fuel quantity is lessthan 1425/1500 pounds).

S NORM-AL

pointer indicates approxi
mately 2675/2810 pounds. FR pointerindicates approximately 3100/3250 ( D 1800/
1890) pounds.

214

When AIM9's are loaded, the PX II
SMS, PX III ST STA switch must be
maintained in the PX II SMS, PX III ST
STA position while taxiing or airborne.
If recycling of the PX II SMS, PX III ST
STA switch is required, power should
not be off longer than 8 seconds at any
one time.

S Communications.
S Steerpoints.
S Profile data (including bingo fuel).
S MFD master mode data.

T.O. GR1F16CJ1*20. MFD's-As desired.

S OVRD lights - On (both cockpits).

21. VHF radio-As desired.

S Selected cockpit stick - Operative.

After FLCS BIT completed:*22. DBU-Check.

*25. MPO - Check.

a. DIGITAL BACKUP switch-BACKUP.
Verify that DBU ON warning lightilluminates.
NOTE
The AOA indicator displays the middle
AOA value selected. The indication
may change in gusty wind conditions.
b. Operate controls-All surfaces respondnormally.
c.

S Other cockpit stick - Inoperative.

DIGITAL BACKUP switch-OFF.
Verify that DBU ON warning light goesoff.

23. Trim-Check.

S TRIM/AP DISC switch-DISC.
S Stick TRIM button-Activate in roll andS Stick - Full forward and hold; notehorizontal tail deflection.

S MPO switch - OVRD and hold.
S Confirm that horizontal tail trailing edges

move farther down.

S Stick and MPO switch - Release.
S Confirm that the horizontal tail returns to

its original position.

*26. Operate controls - All surfaces respondnormally; no FLCS lights on.
*27. AR system (if required) - Check.

S AIR REFUEL switch - OPEN, RDY light on,
DISC light off.

NOTE

pitch.

S No control surface motion.
S No TRIM wheel or indicator motion.

S TRIM/AP DISC switch-NORM.
S Stick TRIM button-Check and center.
S Control surface motion.
S TRIM wheel and indicator motion.

S Rudder trim check.
S YAW TRIM knob - Check and center.

*24.

D

FLCS override - Check.

S DF STICK CONTROL - Selected cockpit.
S DR Stick indicator - As selected.
S Selected cockpit paddle switch - Depress.

If both the RDY and DISC lights are on,then one or more relays are failed. Thiscondition must be corrected prior toAR.

S D A/R DISC button - Depressed. DISC lighton, RDY light off; 3 seconds later, RDY lighton, DISC light off.

S AIR REFUEL switch - CLOSE, RDY lightoff.

*28. Brakes - Check both channels; then return to
CHAN 1.

S One brake pedal - Depress.
S Confirm (ground crew) - Brake activates.

S Opposite brake pedal - Depress.
S Confirm (ground crew) - Initial brake does

not activate.

S Repeat above steps for opposite brake.
215

T.O. GR1F16CJ1
29. Antiice - Check.

33. EPU switch-NORM.

PW 229 Perform the following:

34. Intercom (ground crew) - Disconnect.

S Insure that the engine has been stabilized at

35. Avionic BIT's - As desired.

S TACAN preflight check - Complete.

idle rpm for at least 1 minute with ANTI ICE
switch ON.

S TACANpower knob - CW. (Allow 90
seconds for warmup.)

S ANTI ICE switch - OFF.
S Monitor FTIT when the ANTI ICE switch

S HSI CRS knob - Set COURSE 180.

is moved to OFF. If FTIT does not decrease
by at least 10_C within 15 seconds, abort
the aircraft.

S INSTR MODE knob - TCN.
S Select TACAN BIT on MFD test page.

S ANTI ICE switch - AUTO or ON as required.
GE 129S HSI range and course warning flags - In

Perform the following:view.

S Insure that the engine is stabilized at idle

S HSI bearing pointer may temporarily slew

S ANTI ICE switch - OFF, then ON for 45

S HSI range and course warning flags - Out

to 270 degrees.

rpm.

of view.

seconds.

S HSI range indicator - Indicates 000 ("0.5)

S If ENG A/I FAIL PFL occurs:

nm.

S HSI bearing pointer - Indicates 180 ("3)

S Stabilize engine at 80 percent rpm.

degrees.

S ANTI ICE switch - OFF for 15 seconds,

S HSI CDI - Center within "1/2 dot.

then note FTIT.

S ANTI ICE switch - ON while monitoring

FTIT.

S If FTIT fails to increase a minimum of 10_C

within 30 seconds, abort the aircraft if
flight may encounter areas of known orsuspected icing conditions.

S ANTI ICE switch - AUTO/ON.
30. EPU GEN and EPU PMG lights - Confirm off.
If either light is illuminated, cycle the EPU
switch to OFF, then back to NORM. If either
light stays on, abort the aircraft.
31. EPU switch-OFF.
NOTE
To prevent inadvertently selecting
ON, lift only the aft section of the split
guard.
32. Ground safety pins (ground crew) - Remove.
216

S HSI TOFROM indicator - TO.

36.

C DF Seat - Adjust to design eye.

37.

PX III OBOGS-Check (at least 2 minutes
after engine start).

S OBOGS BIT switch-BIT.
S Verify

that OXY LOW warning lightilluminates for 10 seconds, then goes off.

S Pressure-Check 2540 psi.
S Mode lever-PBG/ON (as required).
S Diluter lever-NORM.
S EMERGENCY lever-NORM.
S FLOW indicator-Check.
S EMERGENCY lever-EMERGENCY.

Check for positive pressure and mask andhose/connector leakage.

S EMERGENCY lever-NORM.

T.O. GR1F16CJ1

BEFORE TAXIS PX II CNI, PX III C & I knob - UFC.

1. Canopy - Close and lock.
The canopy may be partially opened fortaxiing if required for increased visibility.
2. HAVE QUICK radio - Set and check (ifrequired).
NOTE
PX III The HAVE QUICK radio accepts
the first TOD signal that is received
after power up. When the EGI is firstpowered on, the first EGI TOD signal
will likely not be synchronized with the
GPS constellation universal time
coordinated (UTC) time since 3045
seconds are typically required forsatellite acquisition. To insure that the
HAVE QUICK radio is using a
synchronized UTC TOD, another TOD
should be manually received after anadequate GPS track has beenachieved. The best indication of anadequate GPS track is to access the
NAV STATUS DED page (LIST 4) andverify the horizontal accuracy status.
A GPS HIGH estimated horizontal
accuracy status (or less than 300 feet if
the selected accuracy display format isin feet) should ensure that the EGI
TOD signal is synchronized with UTC
time. Alternatively, verify that 4
satellites are being tracked by the EGI
by confirming that a 4 is displayed for
EGI Miscellaneous Parameter (EGIM)
200.

Upfront controls:

S Select the UHF page on DED by depressing
the COM 1 override button.

S PX II CNI, PX III C & I knob - BACKUP.
S Mode knob - PRESET.
S After the MWOD and current date (ifrequired) are entered, select channel 20,
manually set 220.000 MHz, and depress the
PRESET button (failure to do this renders
the UHF radio inoperative).

S Update the TOD by positioning the DED

asterisks around the TOD label on either the
UHF or HAVE QUICK pages and thendepressing the ENTR or MSEL button.
Request a TOD transmission from anotheruser. In either mode, the transmission must
be received within 1 minute of depressing the
ENTR or MSEL button.

S Operating mode - UHF (normal) or HQ (AJ).
S Transmission status - MAIN or BOTH.
S Select an AJ net number with the manualfrequency knob.

S Bandwidth - NB or WB.
Backup controls:

S PX II CNI, PX III C & I knob - BACKUP.
S Function knob - MAIN or BOTH.
S Mode knob - PRESET.
S After the MWOD and current date areentered, select channel 20, manually set
220.000 MHz, and depress the PRESET
button (failure to do this renders the UHFradio inoperative).

S Enter the TOD by requesting a TOD

transmission. To receive TOD update for thenormal mode (UHFTOD), position the PX II
A32T knob, PX III TTONE switch to T andthen return to a normal channel in either
manual frequency or preset channel mode.
To receive TOD update for the AJ mode
(HQTOD), PX II rotate the A32T knob to
T and then back to A, PX III place TTONE
switch to T. A TOD abbreviated time update
can be received on the selected AJ net on
which the TOD is being transmitted. In
either mode, the transmission must be
received within 1 minute of selecting T.

S Select an AJ net number with the manualfrequency knobs.

S PX II A32T, PX III A32 knob-A.

Change 1217

T.O. GR1F16CJ1*3. Altimeterand altitude indications - Set andcheck.
Check that the altimeter readings in
ELECT and PNEU are "75 feet of a known
elevation and are "75 feet of one another.
Also check that the ELECT mode altimeterreading is "75 feet of the altitude displayedin the HUD.
4. Exterior lights-As required.
5.

PX II INS, PX III EGI knob-NAV.

NOTE
Constant rudder pedal pressure
should not be required to maintain astraight ground track during taxi
(based on level surface and no asym
metries or crosswind).
*2. Heading-Check.
*3. Flight instruments-Check
operation.

forproper

BEFORE TAKEOFF
ATF or manual TF is prohibitedfollowing any INS/EGI alignment
other than a full performance NORM
gyrocompass ground alignment withflashing RDY/ALIGN displayed.
6. Chocks (ground crew)-Remove.

If the aircraft has flown in the past 2
hours, do not use the parking brake
except for an emergency. (Does not
apply if drag chute was used.) Parking
brake use may cause residual heatdamage to brakes and may increase theprobability of a subsequent brake fire.

*1. ALT FLAPS switch-NORM.
2. MANUAL TF FLYUP switch-ENABLE.
3. Trim-Check pitch and yaw trim centeredand roll trim as required.

Failure to center yaw trim may result
in severe control difficulties during
takeoff.
4.

C  DF ENG CONT switch-PRI (guard down).

5.

DR ENG CONT switch-NORM (guard down).

6. Speedbrakes-Closed.

TAXI

7. Canopy-Close, lock, light off.

Refer to figure 26 for turning radius and groundclearance.

8. IFF-Set and check.

*1. Brakes and NWS-Check.

FTo minimize heat buildup in the brakeassemblies, do not ride the brakes tocontrol taxi speed. Use one firm
application of the toe brakes to slow
the aircraft.

FHydraulic fluid ignites at relatively

low temperatures. If advised of ahydraulic leak around the brakes,
request firefighting equipment and do
not taxi.

218

9. External tanks (ifinstalled)-Verifyfeeding.
If three tanks are installed, verify that thecenterline tank is feeding. This action
checks that pressurization is available to alltanks.
10. FUEL QTY SEL knob - NORM.
NOTEThe FUEL QTY SEL knob must be in
NORM for operation of the automatic
forward fuel transfer system, trappedfuel warning, and for the BINGO fuelwarning computation to be based on
fuselage fuel.
11. STORES CONFIG switch-As required.

T.O. GR1F16CJ1*12. GND JETT ENABLE switch-As required.
NOTE
Selecting ENABLE starts the DWAT
3minute timer.
*13. Harness, leads, and antig system-Check.

S OXYGEN mode lever-PBG.
S Antig system-Test.

Depress TEST button to check that gsuit
inflates, comfort zippers are closed andremain closed, and hose remains con
nected and that pressure is supplied tomask, helmet bladder, and vest.

Attempting EPU/GEN TEST at idle
rpm results in abnormally low EPU
run speed. Low EPU speed may cause
electrical bus cycling and possible
damage to electrical equipment.
Indications may include blinking EPU
GEN light or audible clicking ofelectrical contactors as the bus cycles.
If electrical bus cycling is experienced,
abort the aircraft and do not taxi.

S EPU/GEN TEST switch-EPU/GEN andhold. Check lights:

S EPU AIR light-On.
S EPU GEN and EPU PMG lights-Off (maycome on momentarily at start of test).

Delayed or slow inflation of the gsuit
could indicate impending antig sys
tem failure. Complete loss of gprotec
tion could occur with no warning.

S FLCS PWR lights-On.
S EPU run light-On for a minimum of 5

seconds.

S OXYGEN mode lever-As required.

FFailure to select PBG when PBG

operation is desired can contribute togray/blackout or loss of consciousness.

FPlacing the OXYGEN mode lever to

PBG without wearing the CRU94/P
harnessmounted connector can expose the pilot to injury producing levelsof oxygen pressure should a failure
occur within the oxygen regulator.

NOTE
If EPU run light does not come on
within 10 seconds after EPU/GEN is
selected, release EPU/GEN TEST
switch, advance the throttle to idle
rpm plus PW 229 10 percent (85percent rpm maximum), GE 129 15
percent (90 percent rpm maximum),and reselect EPU/GEN. If the EPU run
light does not illuminate within 10
seconds or if it cycles on and off, abort
the aircraft.

S EPU/GEN TEST switch-OFF.
S Throttle-IDLE.

14. EPU-Check:
NOTE
FEPU may be checked anytime afterengine start.

S EPU safety pin (ground crew)-Checkremoved.

S Voltage transients may cause fault lightindications. Reset CADC, AVIONICS FAULT
and PW 229 ENGINE FAULT caution lights
as necessary and check MSL ALOW value. A
power cycle of individual systems may also be
required if MFL's occur and remain after the
MFD TEST page is cleared.

S OXYGEN-100%.

S OXYGEN-NORMAL.

S Engine rpm-Increase PW 229 5 percent,

S EPU exhaust (ground crew)-Check for no

GE 12910 percent above normal idle.

airflow.

219

T.O. GR1F16CJ1NOTEFAirflow is indicative of a failed highpressure bleed valve which could
result in an EPU overspeed in flight. If
airflow is detected, abort the aircraft.

FIf required to shut down and restart theengine after completing the EPU checkor if the EPU safety pin is reinstalled forany reason, the EPU check must be
reaccomplished prior to flight to insure
proper EPU operation and closure of thehighpressure bleed valve.

15. FLIR-As required.
16. TFR-As required.
17. PROBE HEAT switch-PROBE HEAT.
Turn probe heat on at least 2 minutes prior
to takeoff anytime icing of probes is possible.
When probe icing is not suspected, delay
selection of probe heat as long as possible
prior to takeoff.

PX III MSL FLOOR value of 0 feet may
indicate a MMC autorestart occurred,and mission planning data has been lost.

*22. All warning and caution lights-Check.
23. Adjustable sliding holder (when utility lightis not in use)- C DF Full forward, rotated cw,
and secured.

Failure to properly secure utility lightand adjustable sliding holder can
result in stick interference.
24. TGP-Stow.

TAKEOFFNORMAL TAKEOFF
Refer to figure 22. When ready for takeoff:

S Advance throttle to approximately PW 229 85,
GE 12990 percent rpm.

S Verify parking brake disengaged.
FProbe internal icing must be suspectedanytime the aircraft has been exposed
to near or below freezing conditions on
the ground. Internal icing may bedifficult to see and may remain present
even when current conditions do notappear conducive to ice formation.

FAOA probes may become hot enough toshut off probe heaters and cause
illumination of the PROBE HEAT
caution light. Refer to CAUTION
LIGHT ANALYSIS, Section III, todetermine if the shutdown was due tooverheat or system failure.

FIf probe heat is on or has been on, heatin probes may be sufficient to cause
injury if touched.

S Check engine instruments for normal indications.
S Release brakes prior to exceeding PW 229 85,
GE 12990 percent rpm.

S Advance throttle to desired thrust.
NOTE
An engine runup check is not required
if conditions require a rolling takeoff.
Maximum FTIT and rpm vary with temperature andpressure altitude but stabilize in 515 seconds.
Normal engine operation during an MIL takeoff isindicated by a exhaust nozzle position of:

S PW 229 20 percent or less.

*18. Ejection safety lever-Arm (down).

S GE 129 15 percent or less after 5 seconds at MIL.

*19. Flight controls-Cycle.

Normal engine operation during an AB takeoff isindicated by:*20. OIL pressure-Check psi.
*21.

PX III ALOW MSL FLOOR Data-Check.

220

S PW 229 The nozzle beginning to open within 5
seconds after selecting AB.

T.O. GR1F16CJ1S GE 129 The nozzle preopening up to 10 percent

more than MIL nozzle position when AB is first
selected. AB lightoff should occur within 5 seconds(greater than 40_F) or 10 seconds (40_F or less)after AB selection and is indicated by increasingfuel flow and nozzle position.

FSince LG and TEF retraction occurs

simultaneously, LG retraction should
not be rushed after takeoff. The
reduction in lift may cause the aircraft
to descend and contact the runway.

NOTE
Spacing of less than 15 seconds
between aircraft when AB is used bypreceding aircraft increases the probability of an AB blowout or no light due
to hot gas ingestion.

FDue to low aft stick forces required for

takeoff, use caution to avoid earlyrotation.

FIf

any FLCS PFL occurs during
takeoff, a WOW switch problem may
be indicated. If the fault clears, themission may be continued. Notify
maintenance of the occurrence after
the flight.

When airspeed is approximately 10 knots below
computed takeoff speed for nonAB or 15 knots for
AB, initiate rotation to establish takeoff attitude(812 degrees). Do not apply aft stick at airspeed
lower than 1015 knots below computed takeoff
speed. Early rotation can lead to overrotation,
skipping, wallowing due to early liftoff, andincreased takeoff distance.

FInsure that the LG handle is placed

fully up. The handle can stop in anintermediate position which retracts
the LG; however, the handle is notlocked and may lower under high g
conditions.

As aircraft lifts off, LEF's extend downward. Retract
LG when safely airborne. Insure LG is up and locked
before exceeding 300 knots. TEF's retract when the
LG handle is raised.

Normal Takeoff (Typical)

TAKE OFF AT
PREDETERMINEDTAKEOFF SPEED

USE NWS UNTIL RUDDER IS
EFFECTIVE (APPROXIMATELY
70 KNOTS)ROTATE AT
PREDETERMINEDROTATION SPEED

GR1F-16CJ-1-0098X37

Figure 22.

221

T.O. GR1F16CJ1TAKEOFF WITH ASYMMETRIC STORES
Roll trim should be set prior to takeoff withasymmetric stores to prevent wing drop. The amountof roll trim required for various asymmetric store
weights is shown in T.O. GR1F16CJ11, Part 2.
NOTE
It is possible to exceed the roll trim
authority of the aircraft for an onspeed
takeoff with a net asymmetric (rolling)
moment less than aircraft takeoff
limits. Refer to ASYMMETRICSTORES LOADING, Section V.
When ARI activates after takeoff, roll trim forasymmetric stores causes a rudder input that cancause aircraft yaw away from the wing with theasymmetric store. This yaw is easily controllable bypilot rudder inputs.

CLIMB
The climb schedules are defined by airspeed/machnumber or mach number only. When airspeed/machnumber is shown, climb at the scheduled airspeed tothe scheduled mach number, then maintain themach number to the desired altitude. When startinga climb at an altitude above the airspeed/machtransition point, climb at the scheduled machnumber.
The recommended MIL climb schedule to optimum
cruise altitude is tabulated versus drag index in T.O.
GR1F16CJ11, Part 3. For MIL, the schedule
provides optimum range performance. The recom
mended MIL climb schedules to altitudes other than
optimum cruise altitude are shown in T.O.
GR1F16CJ11 and T.O. GR1F16CJ1CL1.
The recommended MAX AB climb schedule istabulated versus drag index in T.O. GR1F16CJ11,
Part 3. The MAX AB climb schedule provides
minimum time to climb performance.

222

CLIMB/INFLIGHT/OPERATIONALCHECKSAt frequent intervals, check the aircraft systems,
engine instruments, cockpit pressure, and oxygen
flow indicator and system operation. Monitor fuel ineach internal and external tank to verify that fuelis transferring properly by rotating the FUEL QTY
SEL knob and checking that the sum of the pointers
and totalizer agree and that fuel distribution iscorrect.

Maximum fuel transfer rate is 18,000
pph from the 300gallon fuel tank or
30,000 pph from the 370/600gallon
fuel tanks. Maintaining fuel flow
above these values while the externaltank(s) is feeding results in a decreaseof internal fuel. Prolonged operation
under these conditions may result in
the rapid depletion of fuselage fuel andrender fuel transfer by siphoning
action inoperative. Without siphoning
action, fuel transfer to the fuselage
tanks is provided by the wing turbine
pumps at a maximum rate of 6000 pph.
A fuel flow rate greater than 6000 pph
continues to deplete fuselage fuel.
Under these conditions, the externalfuel tank(s) may appear slow to feed
and a fuel imbalance may result.
Prolonged AB operation in a three tankconfiguration may result in engine
flameout prior to depletion of external
fuel.
1. Fuel-Check quantity/transfer/balance.
2. FUEL QTY SEL knob-NORM.
NOTEThe FUEL QTY SEL knob must be in
NORM for operation of the automatic
forward fuel transfer system, trappedfuel warning, and for the BINGO fuelwarning computation to be based on
fuselage fuel.

T.O. GR1F16CJ1
3. Oxygen system-Check.

*6. Attitude references-Check ADI/HUD/SAI.

4. Cockpit pressurization-Check.
7. ANTI ICE switch-As required.
8. TGP-Stow.
The CABIN PRESS caution light does
not illuminate until cockpit pressure
altitude is above 27,000 feet. At
ambient altitudes where hypoxia ispossible, do not remove the oxygenmask without first checking cockpit
pressure altitude.
5. Engine instruments-Check.

AIR REFUELING PROCEDURES
Refer to T.O. 11C1 for general AR procedures and
to AIR REFUELING PROCEDURES, Section VIII
for specific AR procedures.

DESCENT/BEFORE LANDING
1. Fuel-Check quantity/transfer/balance.
2. Final approach airspeed-Compute.
3. DEFOG lever/cockpit heat-As required.
4. Landing light-On.
*5. Altimeterand altitude indications -Check
altimeter setting, ELECT versus PNEU mode
altimeter readings, and ELECT mode altitude
versus altitude displayed in HUD.
For subsonic flight below 20,000 feet MSL
with vertical velocity less than 500 fpm, thedifference between ELECT and PNEU
mode altitudes should not exceed 270 feetand the difference between the ELECT
mode altitude and the altitude displayed in
the HUD should not exceed 75 feet.

FAn erroneous ELECT mode altitude

can be displayed without a CADC
caution light or a transfer to PNEU
mode.

FAn erroneous altitude can be displayedin the HUD without a CADC caution
light.

LANDINGNORMAL LANDING
Refer to figure 23. Fly initial at 300 knots. At the
break, retard throttle and open speedbrakes as
required. On downwind leg, when airspeed is below
300 knots, lower the LG. During base turn, recheckthe LG down and slow to computed final approach
airspeed to arrive on final at 11 or 13 degrees AOA.
Check speedbrakes open and maintain computed
final approach airspeed/AOA on final. Rate ofdescent decreases slightly when entering groundeffect. Reduce thrust gradually to continue thedescent while applying back stick to reduce sink rate
to the minimum practical. Thrust can be reduced
sooner during an 11degree approach than during a
13degree approach. In either case, maintain amaximum of 13 degrees AOA while reducing sinkrate to the minimum practical.

FPhysically confirm that the LG handle

is fully down. The LG handle may
visually appear to be down when in anintermediate position. An intermediate position may allow LG extension
and/or safe indications; however, theLG handle is not locked and LG
retraction could occur during subse
quent inflight or ground operations.

FFailure of the ANTISKID switch canallow it to be bumped/placed towards
PARKING BRAKE while airborne. A
very small movement out of ANTISKID is sufficient to engage theparking brake. Landing with theparking brake engaged will result in
main tire failures upon touchdown.
Change 1223

T.O. GR1F16CJ1NOTEFFailure to depress the LG handle down

permission button prior to attempting
to lower the LG may result in damage
to the electrical solenoid.

FFailure to reduce sink rate, particu

larly at heavier GW's, may cause a firm
landing and structural damage or
failure of the LG.

F D Use of the paddle switch may cause

pitch and/or roll transients as control
is switched from one cockpit to theother.

FAvoid landing directly on approachend arresting cable to prevent possible
cable strike damage to nozzle, speedbrakes, and ventral fins.

FTo deploy the drag chute after touch

down, lift the guard and switch in one
single motion with the side of the index
finger.

FWhen deploying the drag chute in atwopoint aerodynamic braking atti
tude, the drag chute may contact the
runway.

Use twopoint aerodynamic braking until approxi
mately 100 knots; then fly the nosewheel to runway.
Maximum effective twopoint aerodynamic braking
is achieved at 13 degrees AOA. An AOA less than 11
degrees results in significantly reduced twopoint
aerodynamic braking. Although twopoint aerody
namic braking is effective as low as 80 knots,
runway length and condition should be used todetermine when, after decelerating to 100 knots, tolower the nose to the threepoint attitude.

FHorizontal tail contact with the runway is possible if a large roll input ismade at or near touchdown.

FDeploying the drag chute above 170

knots may result in loss of the chute
canopy.

FDrag chute deployment below 90 knots

may result in improper deployment
and damage to the chute.
NOTE

FThe HUD AOA bracket and AOA

indicator display the correct AOA until
NLG WOW. Therefore, these indica
tions are valid references for aircraft
attitude throughout twopoint aerody
namic braking. After NLG WOW, the
AOA indicator displays zero.

FThe LG warning horn and the TO/LDG

CONFIG warning light are inhibited
at approach airspeed above 190 ("4)
knots.

FAft CG approaches may be character

ized by increased pitch sensitivity
which will be most noticeable uponentering ground effect.

Deploy the drag chute (if desired) immediately after
touchdown. The nose may pitch up or down when
the chute is deployed, but the motion is easily
controlled.
224

FDo not touch down with brake pedals

depressed. A failure in either the
touchdown protection circuitry or an
MLG WOW switch can result in locked
wheels and blown MLG tires.

FUse a maximum of 13 degrees AOA fortwopoint aerodynamic braking. Nozzle,
speedbrakes, and ventral fins may
contact runway if 15degree pitch angle
is exceeded.

FDuring twopoint aerodynamic braking, the speedbrakes (43 degrees orgreater open) may contact the cable.

FDuring the landing phase, large/rapidroll control inputs in reaction toturbulence or wake vortices will cause
temporary retraction of one andsometimes both flaperons. This retrac
tion will decrease lift and may induce
a sink rate beyond the structural limitof the landing gear. During rapid
reversal of roll inputs, both flaperons
might move up to a position that will
illuminate the TO/LDG CONFIG
warning light. Display of ISA FAIL
PFL's is also possible. Be prepared toinitiate a goaround if wake turbu
lence is encountered.

T.O. GR1F16CJ1Normal Landing Pattern (Typical)

LEVEL BREAK
THROTTLE
AS REQUIREDSPEEDBRAKESAS REQUIREDAPPROACH BREAKDOWNWIND LEG
AOA
13 DEGREES
(MAX)
LG
DN
SPEEDBRAKESAS REQUIREDAIRSPEEDTAXI300 KNOTS

NWS
SELECTEDLANDING ROLLDRAG CHUTE
AS REQUIREDSPEEDBRAKESFULLY OPEN
(AFTER NLG WOW)
BRAKESAS
REQUIREDTOUCHDOWN
AOA

13 DEGREES (MAX)
FINAL
AOA
13 DEGREES (MAX)SPEEDBRAKESOPENBASE LEGNOTES:
FINAL APPROACH AIRSPEED/13 DEGREES AOA CROSS-CHECK.

AOA
13 DEGREES (MAX)LG
CHECK DN
SPEEDBRAKESASREQUIREDC PW229 135 GE129 136 KNOTS + 4 KNOTS PER 1000 POUNDS OF FUEL/STORE WEIGHTS.
ADD 8 KNOTS FOR 11 DEGREES AOA APPROACH.
D PW229 137 GE129 138 KNOTS + 4 KNOTS PER 1000 POUNDS OF FUEL/STORE WEIGHTS.
ADD 8 KNOTS FOR 11 DEGREES AOA APPROACH.
THE PRECEDING BASELINE AIRSPEEDS ARE BASED ON THE BASIC OPERATING WEIGHT FROM
T.O. GR1F-16CJ-1-1 PLUS FULL AMMO. ACTUAL FINAL APPROACH AIRSPEED AT 11/13 DEGREES AOA
MAY DIFFER BY +/-5 KNOTS DUE TO VARIATIONS IN AIRCRAFT CG.
GR1F-16CJ-1-0099A37

Figure 23.

225

T.O. GR1F16CJ1
After the nosewheel is on the runway, open the
speedbrakes fully and maintain full aft stick for
maximum threepoint aerodynamic braking andwheel braking effectiveness.

SHORT FIELD LANDING (DRY RUNWAY)
NOTE
The following procedures should be
used anytime stopping distance iscritical, whether due to a long, fast,heavy weight, or short field landing.

FCrossing an arresting cable in a three

point attitude above 90 knots
groundspeed with a centerline store
may cause cable strike.

FDo not move SPD BRK switch to open

until the nosewheel is on runway as
speedbrakes may contact runway.

FUntil WOW, forward stick pressure inexcess of approximately 2 pounds
results in full trailing edge down
deflection of the horizontal tails. This
horizontal tail deflection reduces
wheel braking effectiveness. At highspeeds in the threepoint attitude,
forward stick results in excessive loads
on the NLG which can lead to nose tire
failure and possibly cause structural
failure of the NLG.

Smoothly apply moderate to heavy braking todecelerate to taxi speed. Using less than moderate
braking increases the likelihood of a hot brake(s).
NWS should not be engaged above taxi speed unlessrequired to prevent departure from preparedrunway surface.

NWS malfunctions at any speed may
cause an abrupt turn, tire skidding or
blowout, aircraft tipping, and/or
departure from the prepared surface.

226Change 1

When stopping distance is critical, a normal
approach should be made. Select IDLE at or slightly
before touchdown. Touch down as near as possible
to the end of the runway at 13 degrees AOA. Deploy
the drag chute immediately after touchdown.
Twopoint aerodynamic and wheel braking should
be used with the nose held up at 13 degrees AOA
until the nose falls. Pitch must be held at 13 degrees
AOA if twopoint aerodynamic braking is to beeffective. Maximum effort braking is achieved byusing the wheel brakes in conjunction withtwopoint aerodynamic braking. When the wheel
brakes become effective, the nose automaticallylowers. This occurs soon after brakes are applied.
After the nosewheel is on the runway, maintain fullaft stick, open the speedbrakes fully, and use
maximum wheel braking (antiskid on).
For landing on icy/wet runways, refer to LANDING
IN ICY OR WET CONDITIONS, Section VII.
CROSSWIND LANDING
The recommended technique for landing in acrosswind is to use a wing level crab throughtouchdown. At touchdown, the ARI switches out.
Undesirable yaw transients may occur if roll control
is being applied at this time. After touchdown,
perform twopoint aerodynamic braking using therudder to maintain aircraft track down the runway
and flaperon to prevent wing rise. In crosswinds,
the aircraft may drift downwind due to side loads
imposed by the crosswinds or travel upwind due toinsufficient directional control inputs/availability.
As the airspeed decreases, increasing amounts ofrudder are required to maintain track. Maintaintwopoint aerodynamic braking until approxi
mately 100 knots or until roll or directional control
becomes a problem. As the pitch attitude decreases,
the nose tends to align itself with the ground track.

T.O. GR1F16CJ1
Aft stick and fully opened speedbrakes reduce
stopping distance. Apply brakes after nosewheel ison the runway; however, if stopping distance is afactor, refer to SHORT FIELD LANDING, thissection. With all LG on the runway, maintaindirectional control with rudder, differential braking, and NWS if required.
During landing rollout, the main concerns are wingrise (roll control), weathervaning (directional
control), and downwind drift. Wing rise is controlled
by flaperon into the crosswind. Excessive flaperon
deflection degrades directional control. Use rudder
and differential braking to control ground track,
especially on wet or icy runways. Engage NWS if
required to maintain directional control and toprevent departure from the runway. Excessive
differential braking may result in a hot brakecondition. High rudder pedal force may result in a
yaw transient when NWS is engaged. NLG strut
compression is required to engage NWS but
sustained forward stick may result in fullhorizontal tail deflection which decreases weight on
the MLG and thus reduces wheel braking
effectiveness. NWS engagement may be requiredwith the drag chute deployed to control increased
weathervaning tendencies. However, the nose up
pull of the drag chute may prevent early NWS
engagement.

AFTER LANDING

Do not use parking brake. Use only
chocks, if available, or minimum
possible toe brakes pressure to holdthe aircraft stationary. Parking brakeuse may cause residual heat damage tobrakes and may increase the probabil
ity of a subsequent brake fire.
NOTE
FAvoid heavy braking below 20 knots atlight GW's. Heavy braking during
these conditions may cause both MLG
WOW switches to momentarily go to
the air position, which causes theantiskid system to deactivate the toe
brakes. The WOW switches return to
the ground position after 11.5 seconds, restoring braking capability. If
heavy braking resumes, the cycle may
be repeated.

F C NWS disengagements are possible

when taxiing with CG near theinflight aft limit.

Be prepared to release the drag chuteduring the landing rollout if direction
al control or downwind drifting be
comes a problem.

1. DRAG CHUTE switch - NORM/REL as
required.
NOTE

NOTE
Deploying the drag chute during
twopoint aerodynamic braking with acrosswind may complicate aircraft
control.

Turn aircraft into the wind prior toreleasing drag chute.
2. PROBE HEAT switch-OFF.

TOUCHANDGO LANDING
Perform a normal approach and landing. Aftertouchdown, maintain landing attitude, advance thethrottle, close the speedbrakes, and perform anormal takeoff.

If probe heat is on or has been on, heatin probes may be sufficient to cause
injury if touched.

227

T.O. GR1F16CJ1PRIOR TO ENGINE SHUTDOWN
NOTEProlonged ground operation of probeheat may cause failure of AOA probe
heaters.
3. ECM power-Off.
4. Speedbrakes-Close.
*5. Ejection safety lever-Safe (up).
6. IFF MASTER knob-STBY.
7. IFF M4 CODE switch-HOLD.
8. LANDING TAXI lights-As required.
9. ZEROIZE switch-As required.
NOTE
FIf any FLCS single failures occurred

while airborne, they are reported in
the PFL 2 minutes after WOW. The
FLCS FAULT caution light also illumi
nates.

FIf an FLCS SNGL FAIL PFL occurs andFLCS 049 and 070 MFL's are the only
MFL's present on the MFD test page,perform up to three additional FLCS
BIT's and MFL clear actions. If these
FLCS MFL's clear, no writeup isrequired. If these FLCS MFL's do not
clear, inform maintenance.

10. Canopy handle-Up.
NOTE
FUnlock the canopy to insure that the

canopy seal is deflated before thecanopy is opened.

FIf the canopy handle is placed to up

within 2 minutes of WOW, an FDR 024
MFL is generated.

11. Armament switches-Off, safe, or normal.
228

If a flight control related problem was
experienced during the flight, coordi
nate with maintenance to determine if
the contents of the FLCC fault history
table are desired before shutdown of
FLCS power.
1. EPU safety pin (ground crew)-In.
NOTEInstallation of the EPU safety pinshould be delayed until after engine
shutdown under the following condi
tions:
D The

ground crew recovering theaircraft is not familiar with F16
danger areas.

D The aircraft is being recovered byemergency
response
personnel(landing with activated EPU, hotbrakes, etc.).

Place the EPU switch to OFF prior toengine shutdown if the EPU safety pin
is not installed.
2. EGI/INS-Check.

S Steerpoint of current location-Select.
S Miscellaneous data-Record.
S EGI or INS locations 19 (RER), 20 (CEP),

62 (align events), 64 (NAV events), 66 (VX),
and 67 (VY)-Record.
NOTE

EGI or INS radial error rate (RER)
greater than 3 nm/hour or either
velocity (VX or VY) greater than 5 fps
is considered out of tolerance.
3. MFL-Record (as required).
4. AVTR power switch- PX IIUNTHRD.

OFF,

NOTE
Place the AVTR power switch to OFF atleast 15 seconds prior to engine
shutdown to allow the tape tounthread.

PX IIIT.O. GR1F16CJ1
5.

PX II CNI, PX III C & I knob-BACKUP.

6.

PX II INS, PX III EGI knob-OFF.

NOTE
F PX II Turn INS off at least 10 seconds

prior to engine shutdown to insure that
the INU has adequate time to complete
its shutdown sequence.

F PX III Turn EGI off at least 20 seconds

If either light is illuminated, turn the
MAIN PWR switch to OFF. Insure that
the EPU safety pin remains installed
and notify maintenance.
4. MAIN PWR switch-OFF.

prior to engine shutdown to insure that
the INU has adequate time to complete
its shutdown sequence.

7. Avionics-OFF:S HUD thumbwheels.
S SNSR PWR switches.
S AVIONICS POWER switches.
ENGINE SHUTDOWNNOTEGE 129 Delay placing MAIN PWR
switch to OFF until after engine rpm
decreases through 20 percent. Thisdelay should allow the exhaust nozzle
to remain open and makes it easier for
maintenance to accomplish the post
flight inspection.

5. Oxygen hose, survival kit straps, lapbelt,gsuit hose, and vest hose-Disconnect, stow.

S Stow oxygen connector in bracket on rightsidewall. Insure oxygen hose does not
protrude beyond console edge.

A postshutdown engine tailpipe fire ispossible. Ignition may be indicated by amild bang, followed by smoke, fumes, ora small fire in the combustion/turbine
area. Potentially hazardous inlet andexhaust areas should be avoided within
PW 229 10, GE 129 5 minutes afterengine shutdown. This phenomenon
does not cause damage to the engine or
aircraft. If a postshutdown fire occurs,
the engine may be motored with the
JFS for approximately 1 minute toextinguish the fire. If motoring the JFS
is not possible, the fire extinguishes on
its own within a few minutes.
1. Throttle-OFF.
2.

JFS RUN light-Check.
Notify maintenance if the JFS RUN light isflashing after the throttle is placed to OFF.

After main generator drops off line:
3. EPU GEN and EPU PMG lights-Confirm
off.

S Stow lapbelt and survival kit straps on seatcushion.

S Use both hands to disconnect gsuit hose to

avoid excessive force on the hosetoconsole
connection.

FOnehanded or brute force disconnectsof the gsuit connection will causeinternal damage to the hose at thehosetoconsole connection.

FFailure

to properly stow lapbelt,
survival kit straps, oxygen connector,gsuit hose, and oxygen hose may
cause damage to consoles and to theejection seat during seat adjustment.

6. OXYGEN REGULATOR-OFF and 100%.

229

T.O. GR1F16CJ1
7. Aircraft cocked for scramble-Per local
policies and directives.
FFailure to position the oxygen regula

tor to OFF and 100% may result inparticulate contamination of the regu
lator and subsequent damage.

FTo

avoid damage to the oxygen
regulator, do not pull the knob on theend of the mode lever when moving themode lever from ON to OFF.

7. Canopy-Open.

AIRCRAFT ON QUICK RESPONSE STATUS
If the above actions were not completed prior toscramble, normal preflight procedures should be
used.
1. FLCS power-Check.
With MAIN PWR switch in BATT, verify
that FLCS power is good (FLCS PWR lights)
with the aircraft battery.
2. MAIN PWR switch-MAIN PWR.
3. Engine-Start.

FIf winds exceed 30 knots, open the

canopy only as far as needed to enter/exit
the cockpit. Decreasing the canopy angle
reduces the possibility that the canopy
can be blown past full open.

F C A failure of the canopy actuator

could allow the canopy to fall duringtransit. Keep hands and arms out ofthe path of canopy travel duringopening or closing.

4. Canopy-Close and lock.
5. Instruments-Check.
6. SNSR PWR switches-As required.
7. AVIONICS POWER switches-As required.
8.

NOTE
Prior to aligning the INS/EGI using astored heading alignment, accomplish
a normal alignment, except turn the
PX II INS, PX III EGI knob to OFF fromNORM (prior to power being removed),
and do not move the aircraft.

NOTE
If the internal canopy switch is left inthe up position, the canopy automati
cally opens if closed from the outside.

SCRAMBLE

PX II INS, PX III EGI knob-STOR HDG.

9. FLCS BIT-Accomplish.
10. MFD's-As desired.

PREFLIGHT
Perform the following preflight inspections prior toplacing the aircraft on quick response status:
1. EXTERIOR INSPECTION.

11. SMS-As desired.
*12. HUD/ASHM-As required.
13.

PX II INS, PX III EGI knob-NAV.

2. BEFORE ENTERING COCKPIT.
3. COCKPIT INTERIOR CHECK.
4. BEFORE STARTING ENGINE.
5. STARTING ENGINE.
6. AFTER ENGINE START (include EPU checkbut do not remove MLG ground safety pins).
230Change 1

A full performance NORM gyrocom
pass ground alignment, or PX II
extended interrupted alignment (EIA)with flashing RDY/ALIGN display
should be performed prior to flying any
terrain following mission. If degradedalignment occurs, terrain following is
prohibited.

T.O. GR1F16CJ1
Hot refueling is prohibited with an activated EPU,
hung ordnance, hot brakes, or fuel leakage invicinity of AR receptacle during AR. Safety pins forstores and gun must be installed. In the refuelingarea, use minimum thrust for taxiing.

NOTE
FWhen positioning the PX II INS, PX III

EGI knob from STOR HDG to NAV, do
not hesitate in NORM or the alignment
will be lost.

PRIOR TO HOT PIT ENTRY

FIf time permits, the PX II INS, PX III

EGI knob can be moved to NORM at
any time for a normal alignment.

1. AFTER LANDING checks-Complete.

14. EPU GEN and EPU PMG lights-Confirm
off.
If either light is illuminated, cycle the EPU
switch to OFF, then back to NORM. If either
light stays on, abort the aircraft.

pin

NOTE
F D With the AR switch in OPEN, NWS

can be engaged or disengaged fromeither cockpit regardless of the position of the STICK CONTROL switch
and without using the paddle switch.

If either light is illuminated, the EPU
activates using hydrazine when the
EPU safety pin is removed.
15. EPU-Check (if EPU safety
installed since last EPU check).

2. AIR REFUEL switch-OPEN; RDY light on.

FWith the AR switch in OPEN, thewas

AR/NWS light remains off whenever
NWS is engaged.

*3. TACAN power knob-OFF.
*4. GND JETT ENABLE switch-OFF.

If the EPU safety pin is reinstalled forany reason, the EPU check must be
reaccomplished prior to flight to insure
proper EPU operation.
16. Chocks and safety pins (ground crew) Remove.
*17. Brakes and NWS-Check.
*18. Ejection safety lever-Armed (down).
*19. Flight control surfaces-Cycle.

PRIOR TO HOT REFUELING
Perform the following actions prior to refueling:
1. EPU safety pin (ground crew)-Installed.
*2. Personal equipment leads (except oxygenand communication)-As desired.
3. Canopy-As desired.

20. IFF-As required.

HOT REFUELINGHOT REFUELING PRECAUTIONS
Perform dearming prior to entry to the hot refuelingpit. If suspected hot brakes or other unsafe
condition exists, do not enter refueling area. Follow
ground crew directions into the refueling area andestablish communications with the ground crew. If
a malfunction is suspected, stop refueling.

Insure all cockpit items are secure
prior to opening the canopy. With the
canopy open, the engine is susceptible
to FOD from loose cockpit items.
4. Brake and tire inspection (ground crew) Complete.
5. Intercom with refueling supervisor - Estab
lished.
231

T.O. GR1F16CJ1
DURING HOT REFUELING*1. Be alert for visual or voice signals fromrefueling supervisor.
*2. Terminate refueling if intercom contact islost - Visual signal.
*3. Ground control radio frequency - Monitor.
*4. Insure hands are visible to ground crew.
HOT REFUELING COMPLETE

hydrazine leak or EPU incident occurs on a basewhere no disaster response force or bioenvironmental
support is available, the pilot must insure that theaircraft is isolated and the leak contained. Refer to
ACTIVATED EPU/HYDRAZINE LEAK, Section III.

SUPPLEMENTAL PROCEDURESILS PROCEDURES
1. DED - Verify CNI display.
2. TILS button - Depress and release.

1. AIR REFUEL switch - CLOSE.

3. ILS frequency - Key in and ENTR.

2. EPU GEN and EPU PMG lights - Confirm off.

4. DCS - Position asterisks about selectable
items:

3. EPU switch-OFF.
4. EPU safety pin (ground crew) - Removed.
5. EPU switch-NORM.
6. Intercom (refueling supervisor) - Disconnect.
7. Taxi clear of refueling area and configure
aircraft as required.

QUICK TURNAROUNDPRIOR TO ENGINE SHUTDOWN
1. AFTER LANDING checks - Complete.
2. PRIOR TO ENGINE SHUTDOWN checks Complete.
3. Communication with ground crew - Establish(if required).
4. ENGINE SHUTDOWN checks - Complete.
5. Aircraft setup - IAW local procedures.

SPECIAL PROCEDURESS Key in and enter inbound localizer course.

The inbound localizer course must beentered to obtain correct command steering.

S Depress MSEL button to deselect/selectcommand steering; verify CMD STRG ishighlighted, if selected.

5. HSI - Set inbound localizer course.
6. INSTR MODE knob - ILS/TCN or ILS/NAV.
Verify proper indications/symbology on
ADI, HSI, and HUD.
EXTERIOR INSPECTION
Refer to figure 24 for normal preflight inspection.
DANGER AREAS
Refer to figures 28 and 29.

TURNING RADIUS AND GROUND
CLEARANCE
Refer to figure 210.

EPU HYDRAZINE SUPPORT AT NONF16 BASESGSUIT HOSE ROUTING

At nonF16 bases, the pilot is responsible for theaircraft. Response actions shall be limited to
identification of a hydrazine or EPU problem,
isolation, containment, and minimal dilution withwater (1 part water to 1 part hydrazine). No major
neutralization, maintenance, or hydrazine servicing
capability is planned for transient bases. If a

Refer to figure 211 for the recommended gsuit hoserouting.

232STRANGE FIELD PROCEDURES
Refer to Air Force/Command Guidance.

T.O. GR1F16CJ1Exterior Inspection (Typical)
NOTE:

F.
G.
H.

Check aircraft for loose doors andfasteners, cracks, dents, leaks, andother discrepancies.

I.
J.

EPU EXHAUST PORT-CONDITION.
LEF-CONDITION.
STORES AND PYLONS-SECURE (PREFLIGHT
IAW T.O. GR1F16CJ3411CL1).
NAV AND FORM LIGHTS-CONDITION.
FLAPERON-CONDITION.

AFT FUSELAGE-C
1.

GR1F-16CJ-1-0100X37

NOSE-A
1.

FORWARD FUSELAGE:
A. EXTERNAL CANOPY JETTISON DHANDLES(2)-ACCESS DOORS CLOSED.
B. PITOTSTATIC PROBES (2)-COVERSREMOVED.
C. AOA PROBES (2)-COVERS REMOVED;SLOTSCLEAR;FREEDOM OF MOVEMENT CHECKED;ALIGNMENT CHECKED (ROTATE PROBES FULLYTOWARD FRONT OF AIRCRAFT (CCW ON THE
LEFT;CW ON THE RIGHT) AND VERIFY
BOTTOM SLOTS SLIGHTLY AFT OF 6
O'CLOCK AND TOP SLOTS FORWARD);SET
IN NEUTRAL POSITION (BOTTOM SLOT AT 4
O'CLOCK ON THE RIGHT SIDE AND 8
O'CLOCK ON THE LEFT SIDE).
D. STATIC PORTS (2)-CONDITION.
E. RADOME-SECURE.
F.
ENGINE INLET DUCT-CLEAR.
G. PODS AND PYLONS-SECURE (PREFLIGHT IAW
T.O. GR1F16CJ3411CL1).
H. EPU FIRED INDICATOR-CHECK.
I.
ECS RAM INLET DUCTS-CLEAR.

LEFT WING & CENTER FUSELAGE-D

CENTER FUSELAGE & RIGHT WING-B
1.

2.

TAIL:
A. ADG-CHECK.
B. CSD OIL LEVEL - CHECK.
C. BRAKE/JFS ACCUMULATORS-CHARGED(REFER TO FIGURE 25).
D. HOOK-CONDITION AND PIN FREE TO MOVE.
E. DRAG CHUTE ACCUMULATOR-CHARGED.
F.
VENTRAL FINS, SPEEDBRAKES, HORIZONTALTAILS, AND RUDDER-CONDITION.
G. DRAG CHUTE HOUSING-CONDITION.
H. ENGINE EXHAUST AREA-CONDITION.
I.
NAV AND FORM LIGHTS-CONDITION.
J. VERTICAL TAIL LIGHT-CONDITION.
K. FLCS ACCUMULATORS-CHARGED (REFER TO
FIGURE 27).
L. JFS DOORS-CLOSED.

1.

2.

3.

RIGHT MLG:
A. TIRE, WHEEL, AND STRUT-CONDITION.
B. UPLOCK ROLLER-CHECK.
C. DOOR AND LINKAGE-SECURE.
D. LG SAFETY PIN-INSTALLED.
RIGHT WING:
A. HYDRAZINE LEAK DETECTOR-CHECK.
B. EPU NITROGEN BOTTLE-CHARGED (REFER TO
FIGURE 26).
C. EPU OIL LEVEL - CHECK.
D. HYD SYS A QTY AND ACCUMULA
TOR-CHECK.
E. GUNRNDS COUNTER AND RNDS LIMIT-SET.

4.

LEFT WING:
A. FLAPERON-CONDITION.
B. NAV AND FORM LIGHTS-CONDITION.
C. STORES AND PYLONS-SECURE (PREFLIGHT
IAW T.O. GR1F16CJ3411CL1).
D. LEF-CONDITION.
E. FUEL VENT OUTLET-CLEAR.
F.
HYD SYS B QTY AND ACCUMULA
TOR-CHECK.
LEFT MLG:
A. TIRE, WHEEL, AND STRUT-CONDITION.
B. UPLOCK ROLLER-CHECK.
C. DOOR AND LINKAGE-SECURE.
D. LG SAFETY PIN-INSTALLED.
E. LG PIN CONTAINER-CHECK CONDITION.
FUSELAGE:
A. GUN PORT-CONDITION.
B. IFF-CHECK.
C. AVTR-CHECK.
D. DOOR 2317, ENGINE AND EMS GONOGO
INDICATORS-CHECK.
UNDERSIDE:A. NLG TIRE, WHEEL, AND STRUT-CONDITION.
B. NLG TORQUE ARMS - CONNECTED, PIN
SECURE AND SAFETIED.
C. NLG DOOR AND LINKAGE-SECURE.
D. LANDING AND TAXI LIGHTS-CONDITION.
E. LG/HOOK EMERGENCY PNEUMATIC BOTTLE
PRESSURE-WITHIN PLACARD LIMITS (REFER
TO FIGURE 26).

Figure 24.

233

T.O. GR1F16CJ1Brake/JFS
Accumulators
Pneumatic Servicing
TEMPERATURE°FPRESSURE
PSIG-44 to -36-35 to -27-26 to -18-17 to -9-8 to -1
0 to 89 to 17
18 to 26
27 to 35
36 to 44
45 to 53
54 to 62
63 to 71
72 to 80
81 to 89
90 to 98
99 to 107
108 to 116
117 to 125
126 to 13514751625
15251675
15751725
16251775
16751825
17251875
17751925
18251975
18752025
19252075
19752125
20252175
20752225
21252275
21752325
22252375
22752425
23252475
23752525
24252575

Figure 25.

234Change 1

EPU Nitrogen &Alternate LG/Hook
Bottles Pneumatic
Servicing
TEMPERATURE°FPRESSURE
PSIG

100 and higher
50 to 100
10 to 50-60 to +1032503500
28503250
25002850
20002500

Figure 26.

Drag Chute/FLCS
Accumulators
Pneumatic Servicing
TEMPERATURE°FPRESSURE
PSIG

100 and higher
50 to 100
10 to 50-60 to +10

13001400
12001300
11001200
9501100

Figure 27.

T.O. GR1F16CJ1Danger Areas

ENGINE F110-GE-129

MAX AB THRUST
400

F

150

F TO 365 FT

200

F TO 224 FT

300

F

F

1000

F

25 FT

60 KTS TO 276 FT
1200 KTS

600 KTSVELOCITY
PLANE OFROTATION JFS

020300 KTS

36 KTS TO 512 FT

40
60
80
100
120
140
160
DISTANCE FROM EXHAUST NOZZLEFEETPLANE OF ROTATIONENGINE TURBINE

12

TEMPERATURE
900

100

F
500

F TO 350 FT

8

F
300

WITH ENGINE OPERATING, EAR
PROTECTION SHOULD BE WORNDUE TO HIGH NOISE LEVELS

F

4

0

177 KTS TO 107 FT
950 KTS

25 FT1270 KTS

415 KTSVELOCITY

TIRES

0

10

295 KTS

59 KTS TO 400 FT

592 KTS
20
30
40
50
60
70
DISTANCE FROM EXHAUST NOZZLEWITH HOT BRAKES, AVOID INFLATEDMLG TIRE SIDE AREA WITHIN 300 FEET
FOR 45 MINUTES AFTER AIRCRAFT HAS
STOPPED. IF REQUIRED, APPROACH
FROM THE FRONT OR REAR FOR
FIREFIGHTING PURPOSES ONLY.

80
FEET

IDLE THRUST

12100

300 FT300

F TO 142 FT

F

F
150

15 FT

4

8
100

90

TEMPERATURE
200

10

20
200

180

MIL THRUSTENGINE INLETDANGER AREA

100120 KTS

180 KTS

20

120

F

F

8

4

0
30 KTS

120 KTS90 KTS

300 FT10

36 KTS48 KTSVELOCITY
0

ENGINE INLETDANGER AREA

60 KTS

FEET

2000

F

20
30
40
50
60
70
DISTANCE FROM EXHAUST NOZZLE

FEET

600

F

JET WAKE RADIUS

2900

80
FEET

90

4

FEET

ENGINE INLETDANGER AREA

JET WAKE RADIUS

ENGINE30

F TO 1020 FT

JET WAKE RADIUS

100

TEMPERATURE

8100

GR1F-16CJ-1-0101A37

Figure 28.(Sheet 1)

Change 1235

T.O. GR1F16CJ1Danger Areas

ENGINE F110-GE-129

RADIATION AND TEMPERATURE

UPPER TACANANTENNA

UPPER UHF/IFF
ANTENNANOTES:
Distance from radar disc to forward tip of radome = 5 feet.

60

ECM radiation is in a fanpattern fore and aft.

RADAR

60
ECM PODVHF ANTENNALOWER TACAN ANTENNA
60
RADAR

60
FORWARDRADAR
ALTIMETERANTENNA

LOWER UHF/IFF ANTENNAJFS EXHAUST 1175 F
EPU EXHAUST 1600 FECS EXHAUST 325 F

MINIMUM SAFE DISTANCEFROM ANTENNAS IN FEETOPERATING
TRANSMITTERS

VOLATILE FLUIDSPERSONNEL

UPPER AND LOWER UHF/IFF

1

UPPER AND LOWER TACAN

1

VHF

1

RADAR ALTIMETER

EED1

FIRE CONTROL RADAR

30

120

120

INTERNAL ECM

21

1.5

4

AN/ALQ-119

6

6

AN/ALQ-131

15

15

AN/ALQ-176

6

6

AN/ALQ-184

31

6

AN/ALQ-188

6

6

QRC-80-01

6

6
GR1F-16CJ-1-0102X37

Figure 28.(Sheet 2)236

T.O. GR1F16CJ1Danger Areas

HAZARDOUS NOISE LEVEL AREAS

ENGINE F110-GE-129

MIL THRUST (ESTIMATED)MAX AB THRUST (ESTIMATED)FEET 0NOTE:
1000
Noise levels are
symmetrical and800
may be alteredby surrounding
obstacles.
600
LIMIT 5MINUTESPER DAY
WITH EAR
PROTECTIONFEET 01000

30105 dBA

LIMIT 15MINUTESPER DAY
WITH EAR
PROTECTION

200
90

125
dBA

500

120

600
700
800

115 dBA

800
150

1000
180

IDLE THRUST

250

150

10

115
dBA

120

110 dBA
150

115

20
dBA

40

95 dBA

110

SPL

105
50

100
95
90
85

70

80
50

120
dBA

120

60
100

90

125

30
200

105 dBANOISE EXPOSURE LIMITS

130

FEET
0

300
200

400
120
dBA

1000
180

400

130 dBA
125 dBA
200
300

110
dBA

600

60

500

400

400

100 dBA

600

60

135 dBA200
130 dBA

30

800
700

80

90

0

1
2
3
4
5
6
7
8
DURATIONHOURS EXPOSURE LIMIT PER DAYNOTES:
Attenuation factors and exposure limits are for typicalear muffs and earplugs.
SPL
Sound pressure level.
dBA
Adjusted (human ear response) decibels.
1F-16X-1-4031X

Figure 28.(Sheet 3)

237

T.O. GR1F16CJ1Danger Areas

ENGINE F100-PW-229

MAX AB THRUST

2200

200

F

F TO 470 FT

300 F TO 270 FT
4001955 KTS

237 KTS TO 215 FT
652 KTS
10178 KTS TO 220 FT
948 KTS
119 KTS TO 448 FT
59 KTS TO 850 FT
201540 KTSVELOCITY
PLANE OFROTATION JFS

0

20

40

60

80

100

120

140

160

DISTANCE FROM EXHAUST NOZZLEPLANE OF ROTATIONENGINE TURBINE

180

200

FEET

MIL THRUSTENGINE INLETDANGER AREA

400

500 F

1000

12

100 F TO 378 FTTEMPERATURE
600

F
WITH ENGINE OPERATING, EAR
PROTECTION SHOULD BE WORNDUE TO HIGH NOISE LEVELSFF

200 F TO 110 FT300

F1300 KTS

237 KTS

711 KTSVELOCITY

TIRES

0

10

30

40

50

60

70

DISTANCE FROM EXHAUST NOZZLE

80

90

FEET

12

TEMPERATURE
100

F TO 132 FT

8

300 F

300 FT

400 F

4

150 F

15 FT

200 F

0

148 KTSVELOCITY
0
ENGINE INLETDANGER AREA

4119 KTS

300 FT10

89 KTS
20

30

59 KTS
40

50

30 KTS TO 108 FT
60

70

DISTANCE FROM EXHAUST NOZZLE

80

90

8
100

FEET
GR1F-16CJ-1-1101A37

Figure 29.(Sheet 1)238Change 1

4

8
100

IDLE THRUST
250 F

4

119 KTS TO 155 FT

59 KTS TO 373 FT
20WITH HOT BRAKES, AVOID INFLATEDMLG TIRE SIDE AREA WITHIN 300 FEET
FOR 45 MINUTES AFTER AIRCRAFT HAS
STOPPED. IF REQUIRED, APPROACH
FROM THE FRONT OR REAR FOR
FIREFIGHTING PURPOSES ONLY.

178 KTS TO 118 FT

474 KTS8

0296 KTS
25 FT

10

0

296 KTS356 KTS
25 FT

F

20

FEET

100 F TO 1160 FT

1200 FF

JET WAKE RADIUS

3200

F

FEET

800

JET WAKE RADIUS

ENGINE INLETDANGER AREA

30

F
500 F

FEET

ENGINE600

JET WAKE RADIUSTEMPERATURE

T.O. GR1F16CJ1Danger Areas

ENGINE F100-PW-229

RADIATION AND TEMPERATURE

UPPER TACANANTENNA

UPPER UHF/IFF
ANTENNANOTES:
Distance from radar disc to forward
tip of radome = 5 feet.

60

ECM radiation is in a fanpattern fore and aft.

RADAR

60
ECM PODVHF ANTENNALOWER TACAN ANTENNA60
RADAR
ECM ANTENNA60

CENC AIR-MOTOR
EXHAUST 1200 F

FORWARDRADAR
ALTIMETERANTENNA

LOWER UHF/IFF ANTENNAJFS EXHAUST 1175 F
EPU EXHAUST 1600 FECS EXHAUST 325 F

ECM ANTENNA(EACH SIDE)MINIMUM SAFE DISTANCEFROM ANTENNAS IN FEETOPERATING
TRANSMITTERS

VOLATILE FLUIDSPERSONNEL

UPPER AND LOWER UHF/IFF

1

UPPER AND LOWER TACAN

1

VHF

1

RADAR ALTIMETER
FIRE CONTROL RADAR
INTERNAL ECM

EED

1
30

120

120

21

1.5

4

AN/ALQ-119

6

6

AN/ALQ-131

15

15

AN/ALQ-176

6

6

AN/ALQ-184

31

6

AN/ALQ-188

6

6

QRC-80-01

6

6
GR1F-16CJ-1-1102X37

Figure 29.(Sheet 2)

239

T.O. GR1F16CJ1Danger Areas

HAZARDOUS NOISE LEVEL AREAS

ENGINE F100-PW-229

MIL THRUSTMAX AB THRUSTFEET 01000

NOTE:
Noise levels are
symmetrical andmay be alteredby surrounding
obstacles.
LIMIT 5MINUTESPER DAY
WITH EAR
PROTECTIONFEET 0
1000

30

30

800
700

800
600
400

101 dBA

600

60

500
LIMIT 15MINUTESPER DAY
WITH EAR
PROTECTION

106 dBA

200

60

400
300
200

97 dBA

90

102
dBA

112 dBA200

117 dBA

200
300
117 dBA

400

400
500

120

600

700
800

800
1000
180

107 dBA

150

1000
180

IDLE THRUST
130

0

117
dBA

120
112 dBA150

107 dBA

102 dBANOISE EXPOSURE LIMITS

125

FEET
250

122
dBA

600

112 dBA

102 dBA

90

10

120
20

115

150

40

95 dBA

50

50

100
95

60
100

110
105

SPL

200

dBA

30

90
85

70

80

80
90

0
1
2
3
4
5
6
7
8
DURATIONHOURS EXPOSURE LIMIT PER DAYNOTES:
Attenuation factors and exposure limits are for typicalearmuffs and earplugs.
SPL
Sound pressure level.
dBA
Adjusted (human ear response) decibels.
1F-16X-1-4030X

Figure 29.(Sheet 3)240

T.O. GR1F16CJ1

Turning Radius and Ground Clearance

NOTES:*FEET 39

35

25

18

6

Turn radius can be
reduced to 38 feet byusing inside brake.
Maximum NWS deflection.
All distances areapproximate.
Avoid arresting cable
support donuts whentaxiing with centerline
stores installed.

PIVOT POINT
L MISSILE

L MLG
NLG

PROBE

R MLG
R TAILR MISSILE

With large asymmetries, reduce taxi speed,
avoid sharp turns and, when possible, turntoward heavy wing to prevent possible tipping.

2 FEET*6 INCHES

5 FEET

3 FEET 6 INCHES

3 FEET

5 FEET

16 FEET 10 INCHES

4 FEET 6 INCHES

2 FEET
GR1F-16CJ-1-0104X37

Figure 210.

241

T.O. GR1F16CJ1

G-Suit Hose Routing (Typical)

CANOPY SILL
REFERENCE
LINE

LEFT CONSOLEREFERENCE LINE
(SEAT AT MIDHEIGHT)

LAPBELTTORSO
HARNESSSURVIVAL KITSTRAPG-SUIT HOSE

1F-16X-1-0015X

Figure 211.

242

T.O. GR1F16CJ1

SECTION III
EMERGENCY PROCEDURESTABLE OF CONTENTS

Page

INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . .

33

WARNING AND CAUTION LIGHT
AND PILOT FAULT
LIST ANALYSIS . . . . . . . . . . . . . . . . . . . . . . . .

34

FORMAT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

34

GROUND EMERGENCIES . . . . . . . . . . . . . . . .
Fire/Overheat/Fuel Leak (Ground) . . . . . . . .
Hot Start (Ground) . . . . . . . . . . . . . . . . . . . . . .
Hung Start/No Start PW 229 . . . . . . . . . . . . .
Hung Start/No Start GE 129 . . . . . . . . . . . . .
Engine Autoacceleration (Ground) . . . . . . . .
Antiskid Malfunction (Ground) . . . . . . . . . . .
Brake Failure . . . . . . . . . . . . . . . . . . . . . . . . . . .
Hot Brakes . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Main Generator Failure (Ground) . . . . . . . .
Main and Standby Generator Failure(Ground) . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Emergency Entrance and Crew Rescue . . . .
Emergency Ground Jettison . . . . . . . . . . . . . .
Ground Egress . . . . . . . . . . . . . . . . . . . . . . . . . .
Hot Refueling Emergency . . . . . . . . . . . . . . . .
Activated EPU/Hydrazine Leak . . . . . . . . . .
NWS Failure/Hardover . . . . . . . . . . . . . . . . . .

326
326
326
326
326
326
326
327
328
329

TAKEOFF EMERGENCIES . . . . . . . . . . . . . . . .
Delayed Rotation . . . . . . . . . . . . . . . . . . . . . . . .
Abort . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Engine Malfunction on Takeoff . . . . . . . . . . .
Engine Failure on Takeoff . . . . . . . . . . . . . . .
AB Malfunction on Takeoff . . . . . . . . . . . . . . .
Low Thrust on Takeoff or at Low
Altitude (NonAB) PW 229 . . . . . . . . . . . . .
Low Thrust on Takeoff or at Low
Altitude (NonAB) GE 129 . . . . . . . . . . . . .
Engine Fire on Takeoff . . . . . . . . . . . . . . . . . .
LG Fails To Retract . . . . . . . . . . . . . . . . . . . . .
LG Handle Will Not Raise . . . . . . . . . . . . . . .
Blown Tire on Takeoff . . . . . . . . . . . . . . . . . . .

333
333
333
334
334
335

329
329
329
329
332
332
332

335
336
336
336
336
337

INFLIGHT EMERGENCIES . . . . . . . . . . . . . .
CANOPY Warning Light On . . . . . . . . . . . . .
Canopy Loss/Penetration in Flight . . . . . . . .
Drag Chute Deployed in Flight . . . . . . . . . . .
Cockpit Pressure/Temperature
Malfunction . . . . . . . . . . . . . . . . . . . . . . . . . .
EQUIP HOT Caution Light . . . . . . . . . . . . . .
Ejection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Ejection (Immediate) . . . . . . . . . . . . . . . . . .
Ejection (Time Permitting) . . . . . . . . . . . .
Failure of Canopy To Separate . . . . . . . . .
Ejection Seat Failure . . . . . . . . . . . . . . . . . .
Ditching . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Electrical System Failures . . . . . . . . . . . . . . .
FLCS PMG Failure . . . . . . . . . . . . . . . . . . .
Single Generator Failures(in Flight) . . . . . . . . . . . . . . . . . . . . . . . . .
Main and Standby Generator
Failure (in Flight) . . . . . . . . . . . . . . . . . .
Main, Standby, and EPU Generator
Failure . . . . . . . . . . . . . . . . . . . . . . . . . . . .
EPU Malfunctions . . . . . . . . . . . . . . . . . . . .
Uncommanded EPU Operation . . . . . .
Abnormal EPU Operation . . . . . . . . . . .
FLCS RLY Light . . . . . . . . . . . . . . . . . . . . . .
Partial Electrical Power Loss . . . . . . . . . .
Aircraft Battery Failure . . . . . . . . . . . . . . .
Emergency Power Distribution . . . . . . . . .
Engine Malfunctions PW 229 . . . . . . . . . . . . .
Engine Fire PW 229 . . . . . . . . . . . . . . . . . . .
OVERHEAT Caution Light PW 229 . . . .
Engine Vibrations PW 229 . . . . . . . . . . . . .
Oil System Malfunction PW 229 . . . . . . . .
ENGINE FAULT Caution
Light PW 229 . . . . . . . . . . . . . . . . . . . . . .
SEC Caution Light PW 229 . . . . . . . . . . . .
FTIT Indicator Failure PW 229 . . . . . . . .
Zero RPM/Erroneous RPM
Indication PW 229 . . . . . . . . . . . . . . . . . .
Abnormal Engine Response PW 229 . . . .
Nozzle Failure PW 229 . . . . . . . . . . . . . . . .
AB Blowout/Failure To
Light PW 229 . . . . . . . . . . . . . . . . . . . . . .

338
338
338
338
339
340
340
341
341
341
343
343
343
343
343
344
345
347
347
348
349
349
350
351
361
361
362
363
363
364
364
365
365
365
367
368

31

T.O. GR1F16CJ1Engine Stalls PW 229 . . . . . . . . . . . . . . . . . . . . 368
ABAssociated Engine
Stalls PW 229 . . . . . . . . . . . . . . . . . . . . . . 369
NonAB Engine Stalls PW 229 . . . . . . . . . 369
Engine Stall Recovery PW 229 . . . . . . . . . 369
Inlet Buzz PW 229. . . . . . . . . . . . . . . . . . . . . . 370

Engine Failure or Flameout PW 229 . . . . . . 370
Tower Shaft Failure PW 229 . . . . . . . . . . . 370
Low Altitude Engine Failure or
Flameout PW 229 . . . . . . . . . . . . . . . . . . 371
Airstarts PW 229 . . . . . . . . . . . . . . . . . . . . . . . . 371
High Altitude AirstartConsiderations PW 229 . . . . . . . . . . . . . 378
Low Altitude AirstartConsiderations PW 229 . . . . . . . . . . . . . 379
Airstart Procedures PW 229 . . . . . . . . . . . 379
Engine Malfunctions GE 129 . . . . . . . . . . . . . 381
Engine Fire GE 129 . . . . . . . . . . . . . . . . . . . 381
OVERHEAT Caution Light GE 129. . . . 382

Oil System Malfunction GE 129 . . . . . . . . 383
ENGINE FAULT Caution
Light GE 129 . . . . . . . . . . . . . . . . . . . . . . 384
SEC Caution Light GE 129 . . . . . . . . . . . . 384
FTIT Indicator Failure GE 129. . . . . . . . 385

Zero RPM/Erroneous RPM
Indication GE 129 . . . . . . . . . . . . . . . . . . 385
Abnormal Engine Response GE 129 . . . . 386
AB Blowout/Failure To
Light GE 129 . . . . . . . . . . . . . . . . . . . . . . 387
Engine Stalls GE 129 . . . . . . . . . . . . . . . . . . . . 388
ABAssociated Engine Stalls GE 129. . . 388

NonAB Engine Stalls GE 129. . . . . . . . . 388

Engine Stall Recovery. . . . . . . . . 389

Inlet Buzz GE 129

GE 129. . . . . . . . . . . . . . . . . . . . . . 389

Bird Strike GE 129 . . . . . . . . . . . . . . . . . . . . . . 389Engine Overspeed GE 129. . . . . . . . . . . . . . . 389Engine Failure or Flameout GE 129. . . . . . 389

Tower Shaft Failure GE 129 . . . . . . . . . . . 390
Low Altitude Engine Failure or
Flameout GE 129 . . . . . . . . . . . . . . . . . . 390
Airstarts GE 129 . . . . . . . . . . . . . . . . . . . . . . . . 390
High Altitude AirstartConsiderations GE 129. . . . . . . . . . . . . 395

Low Altitude AirstartConsiderations GE 129. . . . . . . . . . . . . 396

Airstart Procedures GE 129. . . . . . . . . . . 396

32

Flameout Landing . . . . . . . . . . . . . . . . . . . . . . 398
Overhead Approach . . . . . . . . . . . . . . . . . . . 399
StraightIn Approach . . . . . . . . . . . . . . . . 3100
IMC Penetration . . . . . . . . . . . . . . . . . . . . . 3100
Landing Phase . . . . . . . . . . . . . . . . . . . . . . 3105
After Touchdown . . . . . . . . . . . . . . . . . . . . 3105
Flameout Landing Procedures . . . . . . . . 3105
Jettison . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3107
Selective Jettison . . . . . . . . . . . . . . . . . . . . 3107
Emergency Jettison . . . . . . . . . . . . . . . . . . 3108
FLCS Failures . . . . . . . . . . . . . . . . . . . . . . . . . 3108
Air Data Malfunctions . . . . . . . . . . . . . . . 3108
AOA Malfunction . . . . . . . . . . . . . . . . . . . . 3109
CADC Malfunction . . . . . . . . . . . . . . . . . . 3110
Servo Malfunction . . . . . . . . . . . . . . . . . . . 3110
FLCS Electronic Malfunctions . . . . . . . . 3110
FLCS Single Electronic Failure . . . . . 3111
FLCS Dual Electronic Failure . . . . . . 3111
FLCS Temperature Malfunction . . . . . . 3112
Autopilot Malfunctions . . . . . . . . . . . . . . . 3112
DBU ON Warning Light . . . . . . . . . . . . . . 3113
TF FAIL Warning Light . . . . . . . . . . . . . . 3113
LEF Malfunction (Symmetric) . . . . . . . . 3114
LEF Malfunction (Asymmetric) . . . . . . . 3115
Trim Malfunction . . . . . . . . . . . . . . . . . . . . 3116
Stick Interference . . . . . . . . . . . . . . . . . . . . 3116
Fuel Malfunctions . . . . . . . . . . . . . . . . . . . . . . 3117
Fuel Management System PFL . . . . . . . 3117
Fuel Leak . . . . . . . . . . . . . . . . . . . . . . . . . . . 3117
Fuel Low . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3118
Hot Fuel/Oil or Gravity Feed . . . . . . . . . . 3118
Fuel Imbalance . . . . . . . . . . . . . . . . . . . . . . 3119
Trapped External Fuel . . . . . . . . . . . . . . . 3121
Trapped CFT Fuel PX III . . . . . . . . . . . . . 3122
Hydraulic Malfunctions . . . . . . . . . . . . . . . . . 3123
Single Hydraulic Failure . . . . . . . . . . . . . 3123
System A Failure . . . . . . . . . . . . . . . . . . 3123
System B Failure . . . . . . . . . . . . . . . . . . 3123
Dual Hydraulic Failure . . . . . . . . . . . . . . 3124
System B and Generator Failure(PTO Shaft) . . . . . . . . . . . . . . . . . . . . . . . 3125
EGI Failure PX III . . . . . . . . . . . . . . . . . . . . . 3126
INS Failures PX II . . . . . . . . . . . . . . . . . . . . . 3127
Total INS Failure PX II . . . . . . . . . . . . . . . 3128
OutofControl Recovery . . . . . . . . . . . . . . . . 3128
Oxygen Malfunction PX II . . . . . . . . . . . . . . 3130
OBOGS Malfunction PX III . . . . . . . . . . . . . 3131
PBG Malfunction . . . . . . . . . . . . . . . . . . . . . . 3131
Smoke or Fumes . . . . . . . . . . . . . . . . . . . . . . . 3131

T.O. GR1F16CJ1LANDING EMERGENCIES . . . . . . . . . . . . . . .
Straightin Landing . . . . . . . . . . . . . . . . . . . .
Simulated Flameout (SFO) Landing . . . . .
Controllability Check . . . . . . . . . . . . . . . . . . .
Cable Arrestment . . . . . . . . . . . . . . . . . . . . . .
Net Arrestment . . . . . . . . . . . . . . . . . . . . . . . .
Landing With A Blown Tire . . . . . . . . . . . . .
Landing With A Blown Main Gear
Tire . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Landing With A Blown Nose Gear
Tire . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
LG Extension Malfunctions . . . . . . . . . . . . .
LG Handle Will Not Lower . . . . . . . . . . .
LG Fails To Extend . . . . . . . . . . . . . . . . . .
Alternate LG Extension . . . . . . . . . . . . . .
Landing With LG Unsafe/Up . . . . . . . . .
Antiskid Malfunction (Landing) . . . . . . . . .
Asymmetric Stores (Landing) . . . . . . . . . . .
NLG WOW Switch Failure . . . . . . . . . . . . . .
Drag Chute Failure . . . . . . . . . . . . . . . . . . . .

3132
3132
3132
3133
3134
3135
3135
3136
3136
3137
3137
3138
3139
3140
3140
3141
3145
3145

INTRODUCTION
This section covers the operation of the aircraft
during emergency/abnormal conditions. It includes a
discussion of problem indications and corrective
actions as well as procedural steps when applicable.
Adherence to these guidelines insures maximum
safety for the pilot and/or aircraft. The situations
covered are representative of the most probablemalfunctions. However, multiple emergencies, adverse weather, or other factors may require
modification of the recommended procedures. Only
those steps required to correct or manage the problem
should be accomplished. When dealing with emergen
cy/abnormal conditions, it is essential to determine
the most correct course of action by using soundjudgment and a full understanding of the applicablesystem(s). When practical, other concerned agencies(i.e., flight lead, tower, etc.) should be advised of the
problem and intended course of action.
When a voice WARNING or CAUTION message isheard, check cockpit indications; then refer to theappropriate emergency procedure for corrective
action.
When structural damage or any other failure that
may adversely affect aircraft handling characteris
tics is known or suspected, a controllability checkshould be performed.

Certain steps (e.g., MASTER CAUTION reset, ELECCAUTION reset) are intentionally omitted from thenumbered procedures. Pilots are expected to performthese actions without prompting, when warranted.
Three basic rules, which apply to all emergencies, areestablished:
1.

MAINTAIN AIRCRAFT CONTROL.

2.

ANALYZE THE SITUATION AND TAKE
PROPER ACTION.

3.

LAND AS THE SITUATION DICTATES.

The following information provides general landing
guidance:
Land As Soon As Possible
An emergency will be declared. A landing should beaccomplished at the nearest suitable airfield
considering the severity of the emergency, weather
conditions, airfield facilities, lighting, aircraft GW,
and command guidance.
Land As Soon As Practical
Emergency conditions are less urgent and, although
the flight is to be terminated, the degree of theemergency is such that an immediate landing at thenearest suitable airfield may not be necessary.

FThe canopy should remain closed

during all emergencies that could
result in a crash or fire such as crashlandings, aborted takeoffs, and arrest
ments. The protection the canopy
affords far outweighs the isolated riskof entrapment due to a canopy
malfunction or overturn.

FEjection is preferable to sliding into anarrestment cable with the NLG col
lapsed. The cable may slide up over thenose with unpredictable and poten
tially dangerous consequences to any
one in the cockpit(s).

33

T.O. GR1F16CJ1
Fault List procedures. Fault trees show interrela
tionships with examples of problem events andcorrective action.
FIf it appears that the aircraft willdepart a prepared surface above
normal taxi speed during an abortedtakeoff or a landing and goaround is
not possible, eject since breakup of
cockpit structure may occur. Retract
ing the LG to prevent departure from aprepared surface is not recommended
since the MLG will probably not
retract symmetrically.

FIf remaining with an aircraft that willdepart a prepared surface, shut down
the engine, if feasible. This action
reduces the potential for fire, reduces
engine damage, and permits EPU
turnoff if an MLG WOW signal is lost.

WARNING AND CAUTION LIGHT
AND PILOT FAULT LIST ANALYSIS
Refer to figures 31, 32, and 33 for analysis andamplification of warning and caution light and Pilot

34FORMAT
The format of Emergency Procedures differs slightly
between the Checklist and the Flight Manual.
Procedures in the Checklist have been grouped bymalfunction category (engine, electrical, etc.) toprovide maximum inflight utility. In the Flight
Manual, procedures are listed by the phase of flightin which the emergency may occur. In the Checklist,
some procedures are split into two independentsidebyside series of steps and are separated by astraight line; in the Flight Manual, these sidebyside steps appear in a continuous column and can be
identified by repeat numbering of steps following
conditional statements beginning with the word if.
Amplification following procedural steps in the
Flight Manual is repeated in the Checklist under theheading Inoperative Equipment, Other Indications,or Other Considerations. A thorough review of thelayout and content of the Checklist and Flight
Manual is recommended prior to inflight use.

T.O. GR1F16CJ1

Warning Light Analysis
WARNING LIGHT

Engine compartment fire

ENG FIRENO

ENG FIRE

THROTTLE TOMINIMUM PRACTICAL THRUSTWARNING LIGHTCONFIGAIRPROBABLE
MALFUNCTION
IN FIRE DETECTSYSTEM

LAND AS SOON AS POSSIBLENOTES:
Refer to:GROUND EMERGENCIES.
GROUND EGRESS.
ENGINE FIRE.
EJECTION.

ENGINE FIREPERSISTS

CAUSE

TO/LDG

TO/LDG

Refer to ENGINE FIRE

Are there
other signs of fire?
(Visual, FTIT, OVERHEAT
Caution Light, Engine
Response, etc.)
YES

LG Handle
UP(No Malfunction)

CORRECTIVE ACTION/REMARKS

CAUSE

EJECT

CORRECTIVE ACTION/REMARKS

Aircraft configuration wrong for
takeoff or landing

A. For in-flight LG problems, refer to
LANDING EMERGENCIES

Any of the following:
1. LG not down and locked
2. TEF's not fully down

B. If TEF's do not extend, execute landing using normal AOA (approximately
20 knots faster than normal)

BELOW 10,000 FEET
MSL, <190 KNOTS
AND AT LEAST 250
FPM DESCENT

LG/TEF'S NOTDOWN *

INCREASE AIRSPEED, DECREASESINK RATE, OR PLACE LG
HANDLEDN* NOTE: LG warning horn sounds.

CONFIGGND
LG Handle
DN(Malfunction)

TEF'S NOT FULLY
DOWN

AIR
NO

BELOW 10,000 FEET
MSL, <190 KNOTS
AND AT LEAST 250
FPM DESCENTABORT AIRCRAFTONE OR
MORE
GREEN
LIGHTSNOT ONSEE A** NOTES:

Are all
WHEELS down
lights green?

YES

ANY/ALL LG NOTDOWN AND
LOCKED **LG warning hornsounds.
LG handle warning
light illuminates.
3 GREEN
LIGHTS

TEF'S NOT FULLY
DOWN

SEE B

GR1F-16CJ-1-0106X37

Figure 31.(Sheet 1)

35

T.O. GR1F16CJ1

Warning Light Analysis PW229
WARNING LIGHT

CORRECTIVE ACTION/REMARKS

CAUSE

Low pressure in one or both hydraulicCheck pressure indicators to determine syssystems, low engine oil pressure, or low tem and condition
oil pressure switch failure
Refer to:OIL SYSTEM MALFUNCTIONHYDRAULIC MALFUNCTIONSHYD/OILPRESS

HYD/OILPRESSCHECK OIL AND
HYD PRESS
INDICATORS

HYD

OIL

OIL PRESSURENORMALOIL PRESSURE
SWITCH FAILURE

ENGINE OIL
SYSTEM MALFUNCTION, MINIMIZE THROTTLE
MOVEMENT.
LAND AS SOON
AS POSSIBLE

OIL PRESSURE
BELOW 15 PSISYSTEM A
HYD PRESS
INDICATOR ONLY
BELOW 1000 PSIMONITOROIL PRESSURE
INDICATORREFER TO OIL
SYSTEM
MALFUNCTIONSYSTEM A
HYDRAULICFAILUREREFER TOSYSTEM A
FAILURENOSYSTEM B
HYDRAULICFAILUREREFER TOSYSTEM B
FAILURENO

DUAL HYDRAULICFAILUREREFER TO DUALHYDRAULICFAILUREPTO SHAFTFAILUREREFER TO SYSTEM B
AND MAIN GENERATORFAILURE (PTO SHAFT)SYSTEM B
HYD PRESS
INDICATOR ONLY
BELOW 1000 PSIIs EPUrun lighton?
YES
Is ELEC
SYS caution
light on?

YES

BOTH HYD PRESS
INDICATORS BELOW 1000 PSI
WITH EPU SWITCHIN NORM OR ON

TOTAL HYDRAULIC FAILURE(EPU MAY STILL RUN AND
POWER THE EPU GENERATOR,BUT EPU RUN LIGHT DOESNOT ILLUMINATE UNLESSSYSTEM A IS RESTORED)EJECT

GR1F-16CJ-1-1107X37

Figure 31.(Sheet 2)36

T.O. GR1F16CJ1

Warning Light Analysis GE129WARNING LIGHT

CORRECTIVE ACTION/REMARKS

CAUSE

Low pressure in one or both hydraulicCheck pressure indicators to determine syssystems, low engine oil pressure, or low tem and condition
oil pressure switch failure
Refer to:OIL SYSTEM MALFUNCTIONHYDRAULIC MALFUNCTIONSHYD/OILPRESS
HYD/OILMONITOR NOZPOS INDICATORPRESSCHECK OIL AND
HYD PRESS
INDICATORS

HYD

OIL

MONITOR OIL
PRESSURE
INDICATOROIL PRESSURENORMALLOW OIL
PRESSURE
SWITCHFAILUREOIL PRESSURE
BELOW 15 PSI

ENGINE OIL
SYSTEM MALFUNCTION, MINIMIZE THROTTLE
MOVEMENT.
LAND AS SOON
AS POSSIBLE

REFER TO OIL
SYSTEM
MALFUNCTIONSYSTEM A
HYD PRESS
INDICATOR ONLY
BELOW 1000 PSISYSTEM A
HYDRAULICFAILUREREFER TOSYSTEM A
FAILURENOSYSTEM B
HYDRAULICFAILUREREFER TOSYSTEM B
FAILURENO

DUAL HYDRAULICFAILUREREFER TO DUALHYDRAULICFAILUREPTO SHAFTFAILUREREFER TO SYSTEM B
AND GENERATORFAILURE (PTO SHAFT)SYSTEM B
HYD PRESS
INDICATOR ONLY
BELOW 1000 PSIIs EPUrun lighton?
YES
Is ELEC
SYS caution
light on?

YES

BOTH HYD PRESS
INDICATORS BELOW 1000 PSI
WITH EPU SWITCHIN NORM OR ON

TOTAL HYDRAULIC FAILURE(EPU MAY STILL RUN AND
POWER THE EPU GENERATOR,BUT EPU RUN LIGHT DOESNOT ILLUMINATE UNLESSSYSTEM A IS RESTORED)EJECT

GR1F16CJ10107A37

Figure 31.(Sheet 3)

37

T.O. GR1F16CJ1

Warning Light Analysis
WARNING LIGHT

CORRECTIVE ACTION/REMARKS

CAUSE
Canopy hooks or lock not secure

CANOPYUPDOWNWARNING LIGHT

POSITION HANDLE
DOWN

SEE A

CANOPY HOOKSNOT FULLYENGAGEDSEE A

CORRECTIVE ACTION/REMARKS

CAUSE

LG handle warning
light illuminatedPush handle down farther into
locked (outboard) position or ifon the ground, inform maintenance
Refer to CANOPY LOSS/PENETRATION IN FLIGHT

Is canopy
handle up
or down?

CANOPYNOTSECURECANOPYA.

LG or LG door not in positioncommanded by LG handle

A. Refer to LG FAILS TO RETRACT
B. Refer to LG FAILS TO EXTEND

ANY/ALL LG OR
LG DOORS NOTFULLY UP AND
LOCKEDCHECKWHEELSDOWN
LIGHTSSEE A

IN FLIGHT

NONILLUMINATED LIGHT INDICATES
ASSOCIATED LG IS NOT DOWN AND
LOCKED (MAY ALSO HAVE LG WARNING HORN AND TO/LDG CONFIGWARNING LIGHT)

SEE B

GR1F-16CJ-1-0108X37

Figure 31.(Sheet 4)38

T.O. GR1F16CJ1

Warning Light Analysis
WARNING LIGHT
TF FAIL

FLCS

CAUSE

CORRECTIVE ACTION/REMARKSTF system failure

Climb to a safe altitude. Refer to TF FAIL WARNING LIGHT

One or more malfunctions

Note PFL(s) displayed and refer to PILOT
FAULT LIST - FLCS

FLCC operating with backup software
program

Refer to DBU ON WARNING LIGHT

Partial pressure oxygen low

Refer to OBOGS MALFUNCTION PX III , thissectionDBU ONPX IIICANOPYOXY LOW

BIT has detected a fault
Regulator pressure below 5 psi

Flashing WARN symbol

One or more red glareshield warning
lights illuminated

WARN

Flashing TRP FUEL and FUEL warning
symbols

Check for specific illuminated warning lightReset by toggling WARN RESET switch on ICP

A trapped external fuel condition is de
tectedFUEL

Reset FUEL by toggling WARN RESET
switch on ICPCheck fuel tank quantities and refer to
TRAPPED EXTERNAL FUELTRP FUEL

Figure 31.(Sheet 5)

Change 139

T.O. GR1F16CJ1

Warning Light Analysis PW229
WARNING LIGHTENGINE
Check rpm and FTIT indications

ENGINE

CORRECTIVE ACTION/REMARKSCAUSEEngine flameout

A. Place throttle to OFF and initiate airstart.
Refer to AIRSTARTS

Engine alternator failure

B. Refer to ZERO RPM

Engine overtemperature

C. Land as soon as possible

Engine warning system failure or
RPM/FTIT indicator failure

D. Land as soon as practical

CHECK RPM AND
FTIT INDICATIONSIs
RPM
abnormally
low?

YES

Is
RPM
erroneously
zero?

NO

ENGINE
FLAMEOUTSEE A

ENGINE
ALTERNATOR
FAILURE

SEE B

YES
NO

Is
FTIT
abnormally
high?

YES

ENGINE OVERTEMPERATURE

SEE C

ENGINE WARNING SYSTEMFAILURE OR RPM/FTIT
INDICATOR FAILURE

SEE D

NO

GR1F-16CJ-1-1109X37

Figure 31.(Sheet 6)310

T.O. GR1F16CJ1

Warning Light Analysis GE129WARNING LIGHTENGINE
Check rpm and FTIT indications

ENGINE

CORRECTIVE ACTION/REMARKSCAUSEEngine flameout

A. Place throttle to OFF and initiate airstart.
Refer to AIRSTARTS

Engine alternator failure

B. Refer to ZERO RPM

Engine overtemperature

C. Land as soon as possible

Engine warning system failure or
RPM/FTIT indicator failure

D. Land as soon as practical

CHECK RPM AND
FTIT INDICATIONSIs
RPM
abnormally
low?

Is
RPM
erroneously
zero?

YES

NO

ENGINE
FLAMEOUTSEE A

ENGINE
ALTERNATOR
FAILURE

SEE B

YES
NO

Is
FTIT
abnormally
high?

YES

ENGINE OVERTEMPERATURE

SEE C

ENGINE WARNING SYSTEMFAILURE OR RPM/FTIT
INDICATOR FAILURE

SEE D

NO

GR1F-16CJ-1-0109A37

Figure 31.(Sheet 7)

311

T.O. GR1F16CJ1

Caution Light Analysis
CAUTION LIGHT

CAUSE

CORRECTIVE ACTION/REMARKSOne or more caution lights on

Check for specific caution light on caution light
panel

NOTE
The MASTER CAUTION light does notilluminate for the IFF caution light.

Reset MASTER CAUTION light

NOTE
The MASTER CAUTION light does not reset when
the ELEC SYS caution light is illuminated. The
ELEC CAUTION RESET button must be depressedor the electrical malfunction cleared to extin
guish ELEC SYS and MASTER CAUTION lights.

FLCS
FAULT

ENGINEFAULTAVIONICSFAULTSEAT NOT
ARMED

FLCS
FAULT

ENGINEFAULTAVIONICSFAULTSEAT NOT
ARMED

ELEC
SYS

SEC

EQUIP
HOT

NWS
FAIL

ELEC
SYS

SEC

EQUIP
HOT

NWS
FAIL

PROBE HEAT

FUEL/OIL
HOT

RADAR
ALT

ANTI
SKID

PROBE HEAT

FUEL/OIL
HOT

RADAR
ALT

ANTI
SKID

CADC

INLET
ICING

IFF

HOOK

CADC

INLET
ICING

IFF

HOOK

STORESCONFIG

OVERHEATNUCLEAROXY
LOWSTORESCONFIG

OVERHEATNUCLEAROBOGSATF NOTENGAGED

EEC________

CABIN
PRESS

ATF NOTENGAGED

EEC________

CABIN
PRESS

FWD
FUEL LOW

BUC__________________

FWD
FUEL LOW

BUC__________________

AFT
FUEL LOW__________________________

AFT
FUEL LOW__________________________PX II

PX III

Figure 32.(Sheet 1)312

T.O. GR1F16CJ1

Caution Light Analysis
CAUTION LIGHT

CAUSE

CADC

CADC

ENGINE
FAULT

Malfunction of CADC or input(s) to it

A. Refer to CADC MALFUNCTION

Loss of valid mach data to engine

B. Refer to CADC MALFUNCTION

NO

CADC

CORRECTIVE ACTION/REMARKS

MALFUNCTION OF
CADC OR INPUT(S)
TO IT

SEE A

LOSS OF VALID
MACH NUMBERDATA TO ENGINE

SEE B

Is
ENGINE FAULT caution
light on?

YES

GR1F-16CJ-1-0110X37

Figure 32.(Sheet 2)

313

T.O. GR1F16CJ1

Caution Light Analysis
CAUTION LIGHT

CAUSE

CORRECTIVE ACTION/REMARKS

Failure(s) of electrical system*Check ELEC control panel for illuminatedlight(s) and depress ELEC CAUTION RESET
buttonFLCS PMG light (in flight)

Failure of FLCS PMG

A.

Refer to FLCS PMG FAILURE

STBY GEN light

Failure of standby generator

B.

Refer to SINGLE GENERATOR FAILURES

MAIN GEN light

Failure of main generator

C.

Refer to SINGLE GENERATOR FAILURESSTBY GEN and FLCS PMG lights

Failure of standby generator and FLCS
PMG

D.

Refer to SINGLE GENERATOR FAILURES

MAIN GEN, STBY GEN, EPU run, EPU
AIR, and (possibly) EPU HYDRAZN lights

Failure of main and standby generators

E.

Refer to MAIN AND STANDBY GENERATORFAILURE

MAIN GEN, STBY GEN, EPU GEN, and
EPU PMG lights on. EPU run light off

Failure of main, standby, and EPU
generators

F.

Refer to MAIN, STANDBY, AND EPU
GENERATOR FAILURE

MAIN GEN, STBY GEN, EPU GEN, EPU
PMG, FLCS PMG, and ACFT BATT TO
FLCS and/or FLCS BATT lights on. EPU
run light off

Failure of all generators

G.

Refer to MAIN, STANDBY, AND EPU
GENERATOR FAILURE*ACFT BATT FAIL light or ACFT BATT FAIL
and FLCS RLY lights

Failure of aircraft battery or battery
charging system

H.

Refer to AIRCRAFT BATTERY FAILURE

FLCS RLY light

Voltage on one or more of the four FLCC
branches connected to the aircraft battery
is inadequate (below 20V)

I.

Refer to FLCS RLY LIGHT

ELEC
SYS

NOTE:
*During ground operations after engine start, the ACFT BATT FAIL light (and possibly ELEC SYS caution light) may flicker on then off (durationof illumination is not long enough to activate the MASTER CAUTION light). This flickering is the result of a nuisance problem that does
not require corrective action by maintenance and it should not occur in flight.

Figure32. (Sheet 3)314

T.O. GR1F16CJ1

Caution Light Analysis
ELEC
SYS

* CHECK ELEC CONTROLPANEL FOR ILLUMINATED
LIGHT(S) AND DEPRESSELEC CAUTION RESET
BUTTONNOTE:* During ground operations
after engine start, the ACFT
BATT FAIL light (and possibly
ELEC SYS caution light) may
flicker on then off (durationof illumination is not longenough to activate the MASTER CAUTION light). This
flickering is the result of a
nuisance problem that does
not require corrective action
by maintenance and it
should not occur in flight.

FLCS PMG LIGHT(IN FLIGHT)

NONE OF THE FLCS BRANCHES ARE RECEIVING POWER FROM THE FLCS PMG

SEE A

STBY GEN LIGHT

STANDBY GENERATOR FAILURE

SEE B

MAIN GEN LIGHT

MAIN GENERATOR FAILURE
STANDBY GENERATOR ON LINEFLCS PMG OPERATINGSEE CSTBY GEN AND
FLCS PMG LIGHTS

STANDBY GENERATOR AND FLCS PMG
FAILURE

SEE D

MAIN GEN, STBY
GEN, EPU RUN,
AND EPU AIR(AND POSSIBLY
HYDRAZN) LIGHTS

MAIN AND STBY GENERATORS FAILUREEPU IS OPERATING AT PROPER SPEED
FLCS PMG IS OPERATING

SEE E

MAIN GEN, STBY
GEN, EPU GEN,AND EPU PMG
LIGHTS ON. EPU
RUN LIGHT OFF

MAIN GEN FAILURE
STBY GEN FAILUREEPU FAILED OR UNDERSPEED
FLCS PMG OPERATINGAIRCRAFT BATTERY POWERING BATTERY
BUSES

SEE F

MAIN GEN, STBY
GEN, EPU GEN,EPU PMG, FLCS
PMG, AND ACFT
BATT TO FLCS
LIGHTS ON. EPU
RUN LIGHT OFF

MAIN GEN FAILURE
STBY GEN FAILUREEPU FAILED OR UNDERSPEED
FLCS PMG FAILURE
AIRCRAFT BATTERY POWERING FLCS
AIRCRAFT BATTERY POWERING BATTERY
BUSES

ACFT BATT FAIL
LIGHT OR ACFT
BATT FAIL AND
FLCS RLY LIGHTSAIRCRAFT BATTERY SYSTEM FAILURE

SEE H

FLCS RLY LIGHT

VOLTAGE ON ONE OR MORE OF THE
FOUR FLCC BRANCHES CONNECTED TO
THE AIRCRAFT BATTERY IS INADEQUATE(BELOW 20V)

SEE I

SEE G

GR1F-16CJ-1-0111X37

Figure32. (Sheet 4)

315

T.O. GR1F16CJ1

Caution Light Analysis
CAUTION LIGHT
AVIONICSFAULTSEAT NOT
ARMED
HOOK

NWS
FAIL

CAUSE

CORRECTIVE ACTION/REMARKS

Avionic fault(s) detected or MUX commu
nication with FLCS or engine lost

Note PFL(s) displayed on PFLD. Depress C DFFACK, DR FAULT ACK button to acknowledge
fault(s) and reset AVIONICS FAULT caution light.
Perform fault recall(s) to determine if the failure
condition still exists

Ejection safety lever up (system safe)

When desired, rotate ejection safety lever down
(armed)Hook not up and locked

Normal landing, touchdown beyond approachend arresting gear

NWS system failure or loss of NWS electri
cal power

Normal landing, NWS is not available. Refer to
NWS FAILURE/HARDOVERNOTES:
1. Does not illuminate for NWS failure due
to hydraulic system B failure or NLG
strut overextension.
2. Possible indication that AR door cannot
be opened or closed.

Not usedNone

Not used

None

FLCS
FAULT

FLCS fault(s) detected

Note PFL(s) displayed and refer to PILOT FAULT
LIST - FLCS

PW229

FUEL/OIL
HOT

Temperature of fuel to engine excessive.
Oil hot function is inoperative

Refer to HOT FUEL/OIL OR GRAVITY FEEDGE129

FUEL/OIL
HOT

Temperature of fuel to engine excessive orengine oil temperature excessive

Refer to HOT FUEL/OIL OR GRAVITY FEED

EEC

BUC

Forward reservoir contains:
FWD
FUEL LOWCLess
than 400 pounds fuel

Refer to FUEL LOW

DLess
than 250 pounds fuelAft reservoir contains:

AFT
FUEL LOWIFF

CLess
than 250 pounds fuelDLess
than 400 pounds fuel
MODE 4 REPLY sw in OUT (backup mode
only); zeroized or not coded; correct code
not selected (A or B); code does not match
code interrogation; or mode 4 inopera
tive; or RF sw in QUIET or SILENT

Figure 32.(Sheet 5)316

Refer to FUEL LOW

Advisory

T.O. GR1F16CJ1

Caution Light Analysis
CAUTION LIGHT

CAUSE

CORRECTIVE ACTION/REMARKS

INLET
ICING

Ice accumulation detected or system malfunction

If in area of known or suspected icing conditions,
position engine ANTI ICE switch to ON

Malfunction in nuclear control circuitry

Advisory

PW229 Engine operating in SEC ormain fuel pump pressure is low

Refer to SEC CAUTION LIGHT

GE129 Engine operating in SEC

Refer to SEC CAUTION LIGHT

Engine fault(s) detected

Refer to ENGINE FAULT CAUTION LIGHT

STORES CONFIG switch is in incorrect
position or loading category in SMS soft
ware disagrees with actual GP/STORE/
LINE loading category

Verify STORES CONFIG switch is in proper position for aircraft loading category. Refer to T.O.
GR1F16CJ12

Ground:Probe heater failure, monitor
ing system failure, or one/both AOA
probe heaters have shut off to prevent
overheat

Ground:Place PROBE HEAT switch to OFF for
1 minute (caution light goes off when OFF isselected); then reselect PROBE HEAT. If caution
light comes on simultaneously with reselectionof PROBE HEAT, a probe heater or monitoring
system failure has occurred. If caution light does
not come on when PROBE HEAT is reselected,
one/both AOA probe heaters were shut off toprevent overheat

In flight:Probe heater(s) or monitoring
system failure

In flight:Place PROBE HEAT switch to PROBE
HEAT, if required, and avoid areas of known orsuspected icing conditions

NUCLEARSEC

ENGINEFAULT
STORESCONFIGPROBE HEAT

Figure 32.(Sheet 6)

317

T.O. GR1F16CJ1

Caution Light Analysis
CAUTION LIGHT
CABIN
PRESS

CORRECTIVE ACTION/REMARKS

CAUSE

Cockpit pressure altitude above 27,000 A. Refer to COCKPIT PRESSURE/TEMPERfeet
ATURE MALFUNCTIONBELOW 27,000
FEET

MALFUNCTION
IN INDICATORSYSTEMCHECK COCKPITPRESSURE
ALTITUDE

CABIN
PRESS

COCKPIT PRESSURE
EXCEEDS 27,000
FEET BECAUSE OF
PRESSURIZATIONFAILUREYESABOVE 27,000
FEET

CONTINUE
FLIGHT

SEE A

Is
AIR SOURCE knob
in NORM?
NOAIR SOURCE
KNOB
NORM

CAUTION LIGHT

CAUSE

CORRECTIVE ACTION/REMARKS

EQUIP
HOT

Avionic equipment cooling air temperature/pressure insufficient

A. Refer to EQUIP HOT CAUTION LIGHT

EQUIP
HOT
NO
ILLUMINATES WHEN
AVIONIC EQUIPMENT
COOLING AIR TEMPERATURE/PRESSURE
IS INSUFFICIENT

AIR SOURCE
KNOB
NORM

Is
AIR SOURCE knob
in NORM?
YES

DEGRADED EQUIPMENT PERFORMANCE
AND/OR DAMAGECAN RESULTSEE A

GR1F-16CJ-1-0112X37

Figure 32.(Sheet 7)318

T.O. GR1F16CJ1

Caution Light Analysis
CAUTION LIGHT

CAUSE

CORRECTIVE ACTION/REMARKSOVERHEAT

Overheat condition in engine compartment, ECS or EPU bays, or MLG wheel
wells

A. Refer to OVERHEAT CAUTION LIGHT

OVERHEATTHROTTLE TOMINIMUM PRACTICAL THRUSTCHECK FIRE
DETECT CIRCUIT

Did
OVERHEAT
caution light
go off?
NO

YES

Did
overheatdetect circuit
check good?

YES

Did
overheatdetect circuit
check good?

YES

OVERHEATCONDITION
IN ENGINE
COMPARTMENTSEE A

OVERHEATCONDITION
IN EPU BAY

SEE A

NO

Is
EPU
running?

YES
TURN EPU OFF
IF IT WAS MANUALLY ACTIVATEDNO

Did
OVERHEAT
caution light
go off?
NO

YES

NO

REACTIVATEEPUOXYGEN 100%

CONT

GR1F-16CJ-1-0113X37

Figure 32.(Sheet 8)

319

T.O. GR1F16CJ1

Caution Light Analysis
CONT
AIR SOURCE
KNOB
OFF

Did
OVERHEAT
caution light
go off?

Didoverheatdetect circuit
check good?

YES

YES

OVERHEATCONDITION
IN ECS BAY

SEE A

NO
NO
LG

OVERHEAT CONDITION
IN MLG WHEEL
WELLS

DN

LAND AS SOON
AS POSSIBLE

YES
Did
OVERHEAT
caution light
go off?

Didoverheatdetect circuit
check good?

YES

NO

OVERHEAT CONDITION IN ENGINE COMPARTMENT, ECS OR EPU BAYS, OR MLG
WHEEL WELLS STILL PRESENTNOOVERHEAT CONDITION IN ENGINE COMPARTMENT, ECS OR EPU BAYS, OR MLG
WHEEL WELLS STILL PRESENTCAUTION LIGHT
ANTI
SKID

SEE A

ANTI-SKID switch OFF or system
malfunction

YES

ANTI
SKID

CORRECTIVE ACTION/REMARKSCAUSEANTISKID
MALFUNCTION

A. Refer to ANTISKID MALFUNCTION

SEE A

Is switch
in ANTI-SKID
position?

NO

POSITION TO
ANTI-SKID
GR1F-16CJ-1-0114A37

Figure 32.(Sheet 9)320

T.O. GR1F16CJ1

Caution Light Analysis
CAUTION LIGHT

CORRECTIVE ACTION/REMARKSCAUSE
NVP or FLCS cannot support ATF

ATF NOTENGAGED

If in ATF, climb to a safe altitude andverify:AIR REFUEL switch CLOSE
ALT FLAPS switch NORM
TRIM/AP DISC switch NORMNo CADC failures
NOTE
Deselect ATF until the cause of thecaution light illumination can be
determined.

RADAR
ALT

II

OXY
LOW

PX III

OBOGS

PX

Malfunction of radar altimeter

Move RDR ALT switch to OFF

Oxygen pressure is below 42 psi or
quantity is below 0.5 liter

A. Descend to 10,000 feet cockpit pressure altitude. Activate emergency
oxygen if required

The ECS air supply for OBOGS has
dropped below 10 psi

Oxygen production stops. BOS provides
oxygen for 35 minutes. Expect OXY LOW
warning light if regulator pressure drops
below 5 psi. Attempt to increase ECS airpressure by increasing throttle setting,increasing airspeed, and/or decreasing
altitudeNOTE
The OBOGS caution light may illuminate asa result of ECS cycling or temporary ECS
shutdown. This is normal as long as the
OXY LOW warning light does not illuminate.

SEE A

YES

PX

II

OXY
LOW

Is oxygen
quantity below
0.5 liter?
NO

YES

Is oxygenpressure below
42 psi?

NO

QUANTITY
INDICATION
MALFUNCTION

GR1F-16CJ-1-0115A37

Figure 32.(Sheet 10)

321

T.O. GR1F16CJ1

Pilot Fault List-Engine
FAULT

PW 229CAUSE

CORRECTIVE ACTION/REMARKSENG A/B FAIL and
ENG THST LOWEngine hardware deterioration/detected
performance loss

Reduce engine rpm to 85% or less, unless required to sustain flight. High thrust levels may
result in further deterioration/performance loss.
Land as soon as possible

ENG A/I TEMP

Antiice valve failed open and/or bleed airtemperature greater than 850°F

Reduce throttle setting to midrange unlessrequired to sustain flight. Operating the engine
ice system failed on
above midrange with anti
antiice
may result in engine stall. Land as soon as practical

ENG A/I FAIL

Engine antiice valve failed in closed
position

Avoid areas of known or suspected icing
conditions

ENG MACH FAIL

The CADC supplied mach number to the
DEEC is no longer available

Supersonic stall protection is inoperative. Do notretard throttle below MIL while supersonic. If
CADC caution light is also on, refer to CADC
MALFUNCTION, this section

ENG A/B FAIL

AB system failure detected

AB RESET switch-AB RESET. Land as soon aspractical if fault does not clear. AB operation ispartially or fully inhibited

ENG THST LOW

Loss of redundant FTIT signals received by
DEEC
DEEC has detected a failed open or miss
ing nozzle

MIL rpm is reduced 7 percent by DEEC. Land assoon as practical
If a failed open or missing nozzle is suspected,
refer to NOZZLE FAILURE, this section

ENG BUS FAIL

Communication lost between EDU andMUX busIlluminates AVIONICS FAULT caution light. A
subsequent engine fault causes a nonresettable
ENGINE FAULT caution light and is not displayedon the PFLDENG PFL DGRD

Communication lost between EDU andDEEC

Do not retard throttle below MIL while supersonic. Only ENG A/I TEMP PFL can subsequently
be displayed

NOTE:
A short duration fault condition may cause display of a PFL without illumination of the ENGINE FAULT caution light.

Figure 33.(Sheet 1)322

T.O. GR1F16CJ1

Pilot Fault List-Engine
FAULT

GE 129CAUSE

CORRECTIVE ACTION/REMARKSENG LUBE LOW

Oil quantity below preset limit

Refer to OIL SYSTEM MALFUNCTION, thissection

ENG A/I FAIL

Engine antiice valve failed in closed position or indication malfunction

Avoid areas of known or suspected icing condi
tions

ENG MACH FAIL

The CADC supplied mach number to the
DEC is no longer available

Supersonic stall protection is inoperative. Do notretard throttle below MIL while supersonic. If
CADC caution light is also on, refer to CADC
MALFUNCTION, this section

ENG BUS FAIL

Communication between engine and
MUX bus lost

Illuminates AVIONICS FAULT caution light.
Other engine PFL's cannot be displayed

ENG A/B FAIL

AB system failure detected

Land as soon as practical if fault does not clear.
AB operation inhibited. If nozzle remains closed
at idle below 0.5 mach, refer to ABNORMAL EN
GINE RESPONSE, this section

ENG EMS FAIL

D Data transmission from DEC lost

Other PFL's either cannot be displayed or, if dis
played, are not reliable

D BIT/selftest received from DEC
indicates a failure
D EMSC BIT/selftest detects a failure
ENG HYB MODE

A PRI fuel flow scheduling problem was
detected

Supersonic stall protection is inoperative. Do notretard throttle below MIL while supersonic.
Check engine response to throttle movement
when subsonic. If engine responds normally,land as soon as practical. If engine does not re
spond normally, refer to ABNORMAL ENGINERESPONSE, this section

NOTE:
A short duration fault condition may cause display of a PFL without illumination of the ENGINE FAULT caution light.

Figure 33.(Sheet 2)

323

T.O. GR1F16CJ1

Pilot Fault List - FLCS
(FLCS warning light illuminated)
FAULT

CAUSE

CORRECTIVE ACTION/REMARKSFLCS AOA WARN

Dual AOA failure

Refer to AOA MALFUNCTIONSFLCS DUAL FAIL

Dual electronic, sensor, or power failure in
one or more axes

Refer to FLCS DUAL ELECTRONIC FAILURE

FLCS LEF LOCKLEF's are locked due to multiple failures,
LE FLAPS switch position, or asymmetry

Refer to LEF MALFUNCTIONSSTBY GAIN

Dual air data failure

Refer to AIR DATA MALFUNCTIONSFLCS BIT FAIL

FLCS BIT has detected a failure

Perform a second FLCS BIT. If fault does not clear,
notify maintenance. Fault only occurs on ground(TF FAIL warning light illuminated)
FAULT

CAUSE

SWIM NVP FAIL

NVP data bad, AMUX wraparound failure,
NVP selfmode failure, or cyclic test prob
lem monitor failure

SWIM RALT FAIL

SDC monitor failure or CARA data bad

SWIM SCP FAIL

Below set clearance failure

SWIM ATTD FAIL

INS attitude estimator failure

SWIM ATF FAIL

NVP ATF select failure

SWIM VEL FAIL

GPS/INS failure

Figure 33.(Sheet 3)324CORRECTIVE ACTION/REMARKS
Refer to TF FAIL WARNING LIGHT

T.O. GR1F16CJ1

Pilot Fault List - FLCS

(FLCS FAULT caution light illuminated for all except FLCS BUS FAIL)
FAULT

CAUSE

CORRECTIVE ACTION/REMARKSFLCS ADC FAIL

First failure of triplex air data input
signalRefer to AIR DATA MALFUNCTIONSFLCS AOA FAIL

First failure of triplex AOA input signal

Refer to AOA MALFUNCTIONSFLCS AOS FAIL

AOS feedback function is inoperative due
to failure

Perform FLCS reset to attempt to clear fault;
fault cannot be reset if INS or CADC is failed
If fault does not clear, the autopilot cannot be
engaged. Position the STORES CONFIG switch
to CAT III if the aircraft is configured with a
33 GP/STORE/LINE loading. Refer to T.O.
GR1F16CJ12 *FLCS FLUP OFF

MANUAL TF FLYUP switch moved to DISABLE

Position the MANUAL TF FLYUP switch as required. A FLCS reset extinguishes FLCS FAULT
caution light

FLCS BIT detects MANUAL TF FLYUP switch
in DISABLE

Position MANUAL TF FLYUP switch to ENABLE.
Rerun FLCS BIT

FLCS A/P DEGR

Autopilot operating outside of attitude
limits or unable to hold commanded
mode

Refer to AUTOPILOT MALFUNCTIONSFLCS A/P FAIL

Autopilot has disconnected or cannot be
engaged due to loss of needed data

Refer to AUTOPILOT MALFUNCTIONSFLCS BUS FAIL

Communication lost between FLCC andMUX busIlluminates AVIONICS FAULT caution light.
Other FLCS PFL's may not be displayed on the
PFLD. Refer to FLCS page on MFD for FLCS PFL's

BRK PWR DEGR

Power supply failure detected in one ormore branches

Refer to FLCS SINGLE ELECTRONIC FAILURE

FLCS CCM FAIL

Erroneous output command detected by
CCM

Refer to FLCS SINGLE ELECTRONIC FAILURE

FLCS HOT TEMPFLCC sensors detect two branches inexcess of 75°C

Refer to FLCS TEMPERATURE MALFUNCTIONSISA ALL FAIL

Two or more ISA's have reported a first
servo valve failure

Refer to SERVO MALFUNCTIONSISA LHT FAIL
ISA RHT FAIL
ISA LF FAIL
ISA RF FAIL
ISA RUD FAIL

Indicated ISA has reported a first servo
valve failure

Refer to SERVO MALFUNCTIONSFLCS SNGL FAIL

Indicates single electronic or sensor fail
ure in one or more axes

Notify maintenance. Fault only occurs on
groundFLCS MUX DEGR

BIT detected degradation of FLCC MUXinterface

FLCS reset will not clear fault. Perform a second
FLCS BIT. If fault does not clear and no otherfaults are reported, the system redundancy isadequate for flight. Notify maintenance after
flight. Fault only occurs on ground

NOTE:* The potential for a departure from controlled flight is significantly increased if the AOS feedback function is inoperative and maneuveringwith 33 GP/STORE/LINE loadings occurs with the STORES CONFIG switch in CAT I.

Figure 33.(Sheet 4)

325

T.O. GR1F16CJ1GROUND EMERGENCIESENGINE AUTOACCELERATION (GROUND)

FIRE/OVERHEAT/FUEL LEAK (GROUND)

If the engine autoaccelerates on the ground, primaryconsideration should be given to shutting the engine
down as quickly as possible. With engine shut down,
only the brake/JFS accumulators are available tosupply hydraulic pressure for braking. Stop theaircraft by making one steady brake application.
When the aircraft is fully stopped, have chocks
installed or engage parking brake. Leave the batteryon until chocks are installed.

An engine or JFS fire/overheat can be detected byflames, smoke, explosion, signal from ground crew,or radio call. FTIT may exceed PW 229 800_C,
GE 129 935_C and, if ac power is available, ENG
FIRE warning or OVERHEAT caution light may
illuminate.
1.

Throttle-OFF.

2.

JFS switch-OFF.

3.

FUEL MASTER switch-OFF.

4.

ENG FEED knob-OFF (if external power
applied).

If fire continues:
5.

Abandon aircraft.

HOT START (GROUND)PW 229 Hot start-FTIT over 800_C. During engine

start, if the FTIT increases at an abnormally rapid
rate through 750_C, a hot start can be anticipated.
GE 129 Hot start-FTIT over 935_C. During engine
start, if the FTIT increases through 750_C while
engine rpm is less than 40 percent, a hot start can be
anticipated.

1.

Throttle-OFF.

2.

FTIT indicator-Monitor.

If FTIT remains above 500_C:
3.

JFS switch-START 2.
Motor engine with JFS until FTIT reaches
200_C or for four minutes (JFS ground
operating limit), whichever occurs first.

HUNG START/NO START PW 229
Hung start-RPM has stopped increasing below
IDLE and FTIT is stabilized at less than 800_C.
No start-Lightoff does not occur within 20 seconds.
1.

Throttle-OFF. Notify maintenance.

HUNG START/NO START GE 129
Hung start-RPM has stopped increasing below
IDLE and FTIT is stabilized at less than 935_C.
No start-Lightoff does not occur within 10 seconds.
1.

Throttle-OFF. Notify maintenance.

326Change 1

1.

Throttle-OFF.

2.

FUEL MASTER switch-OFF.

ANTISKID MALFUNCTION (GROUND)
b2t If a failure affecting braking performance is

detected while the aircraft is moving above 5 knots,
the ANTI SKID caution light illuminates. In mostcases this represents the loss of a wheel speed sensor
signal, and the system switches to an alternate
braking mode. In this mode, if differential braking isapplied (15 percent or greater difference between
pedals), both brakes oscillate between pressure asmetered and no pressure. Braking effectiveness is
reduced by 50 percent or greater. If brake pedals arewithin 15 percent, the system uses the informationfrom the remaining good wheel speed sensor andstopping distance is increased by approximately 25
percent on both wet and dry runways. An ANTISKID caution light which only illuminates below 5
knots indicates a malfunction that does not affectbraking performance. Normal braking and antiskid
are available; however, system redundancy may
have been lost.
Below 20 knots ground speed the alternate
braking mode is less effective. Place BRAKES
channel switch to CHAN 2 and ANTISKID switch toOFF. Braking will be manual.
b2t

LESS b2t In case of an antiskid failure, the ANTISKID caution light illuminates and the brake system
automatically switches to pulsating pressure. In thismode, braking effectiveness is reduced approxi
mately 50 percent; however, in most cases, braking
effectiveness is as good as can be obtained with
ANTISKID switch in OFF while avoiding wheel
lockup and its associated risk of control difficulty.
Short field landing distances are increased approxi
mately 60 percent for dry runway and 25 percent for
wet runway from those normally computed.

T.O. GR1F16CJ1
If the ANTI SKID caution light illuminates (with the
ANTISKID switch in ANTISKID):
1.

DRAG CHUTE switch-DEPLOY (if required).

2.

Brakes-Apply as needed.
NOTE
F b2t Use of maximum symmetric pedal

pressure provides the best stopping
performance. Differential brake only
when essential for directional control.
If the ANTI SKID caution lightilluminated
above
5
knots
groundspeed, the aircraft may oscillate
due to pulsating brake pressure (if 15
percent or greater differential pedal
pressure is applied). Changing brakechannels may restore normal braking.

FLESS b2t Maximum pedal pressure isrequired to obtain approximately 50
percent of normal braking force. If lessthan maximum pedal pressure is used,braking is extremely degraded. The
aircraft will oscillate due to pulsatingbrake pressure. Changing brake channels will not restore normal braking
since the same antiskid signal is usedin both brake channels.

3.

NWS-Engage (if required).

If manual braking is desired or after aircraft is
stopped:4.

b2t BRAKES channel switch-CHAN 2.

5.

ANTISKID switch-OFF.

No antiskid protection is available
with the ANTISKID switch in OFF
b2t and BRAKES channel switch in
CHAN 2. Brakes should be appliedwith caution to avoid wheel lockup andblown tires.

BRAKE FAILURE
Malfunctions in systems which affect normal braking
are described in the emergency procedure which
addresses each specific system. One of the brake
failure modes is the loss of one brake circuit. With thisfailure, both brakes are still available; however,significantly more pedal force than normal isrequired to achieve a specific braking effectiveness.
Another failure mode is loss of brakes on one or both
MLG. Changing brake channels may return thesystem to normal operation. Turning the ANTISKID switch to OFF b2t and confirming BRAKES
channel switch in CHAN 2 may also restore braking;however, the system reverts to manual control andantiskid protection is lost. (Status of the powersource for toe brake transducers can be determinedby testing the FLCS PWR lights on the TEST switch
panel.) Release brake pedal pressure beforechanging channels or turning off the ANTISKID
switch to avoid immediate brake lockup if braking
returns. When moving the ANTISKID switch, bevery careful not to select the PARKING BRAKE
unless that is what is intended. If directional control
is a problem (such as with one brake inoperative on
landing roll), do not hesitate to use NWS. If
conditions permit, consider a goaround if the brakes
are found to be inoperative on landing. Lower hook
if a cable is available. If normal brakes cannot be
restored, do not hesitate to use the parking brake if
a cable is not available. The lower the groundspeed,
the less chance there is for aircraft damage whenusing the parking brake. If the aircraft isaccelerating, use the parking brake early. It may bepossible to cycle the parking brake on and off andstop the aircraft; however, regardless of technique,use of the parking brake may result in blown tires.
Another failure mode is a hydraulic leak in the brakeitself, which might not be apparent until after
twopoint aerodynamic braking. In this case, if acable is not available, the aircraft should be stopped
using the good brake and NWS for directional
control. Once the aircraft is stopped, do not engage
the parking brake; use continuous pedal pressure on
the good brake only. Failure to do so could deplete thehydraulic system and result in total brake failure
prior to chock installation.
Accomplish as many steps as required:

NOTE

NOTE

LESS b2t Below normal taxi speed,
pulsating braking is only marginallyeffective. Stopping distance may be
shortened with antiskid off.

If conditions permit, consider a go
around if the brakes are found to be
inoperative on landing. An approachend cable arrestment is recommended.
327

T.O. GR1F16CJ1
1.

BRAKES channel switch-Change channels.

Perform the following after any event that may resultin hot brakes:
1.

Release brakes prior to changing
brake channels or turning antiskid off.
2.

b2t BRAKES channel switch-CHAN 2.

3.

ANTISKID switch-OFF.

Request firefighting equipment and proceed
directly to the designated hot brake area ornearest area clear of other aircraft andpersonnel.

FIf a hot brake condition is a result of a

dragging brake, taxiing the aircraft
worsens the condition.

Release brakes prior to changing
brake channels or turning antiskid off.
4.

NWS-Engage (if required).

5.

HOOK switch-DN.

FAny leaking hydraulic fluid may be

ignited by hot wheel and brakesurfaces.

FWheel fusible plugs may relieve tirepressure at anytime during the 15
minutes after brake application.

If arresting cable is not available or if at low
groundspeed:
6.

FWith hot brakes, avoid inflated MLG

tire side area within 300 feet for 45
minutes after aircraft has stopped. If
required, approach from the front orrear for firefighting purposes only.

ANTISKID switch-Intermittent PARKINGBRAKE, then ANTISKID.

When in the hot brake area:
If in a congested area, use the parking
brake immediately to stop.

2.

Align aircraft with nose into wind if possible.

HOT BRAKES
The pilot has the responsibility to determine when a
hot brake condition exists. The pilot evaluates thesituation by analyzing the variables that influence
brake temperature: GW, pressure altitude, OAT,
speed at brake application, etc. Refer to T.O.
GR1F16CJ11, PART 2, BRAKE ENERGY LIMITS- MAXIMUM EFFORT BRAKING. Observations byground crewmembers should also be used as certain
malfunctions that result in overheated brakes, suchas dragging brakes, may not be readily apparent to
the pilot. Perform hot brake procedures anytime hot
brakes are suspected.

FDo not use the parking brake.
FIf battery power is not available, toe

brakes will be inoperative after engine
shutdown.

FDo not turn MAIN PWR switch to OFF

until the nosewheel is chocked.

FAttempt to park in a level area tominimize risk of aircraft rolling if thebrakes should fail after shutdown.

It is impossible for the ground crew to avoid the hot
brake and engine intake danger areas while pinning
the EPU or chocking the aircraft. Therefore, if
conditions permit, the aircraft should be shut down
without pinning the EPU or chocking the wheels.
Release brake pressure as soon as possible tominimize heat transfer between the brake surfaces
and the wheel. This action also relieves hydraulicpressure to the brakes, which if leaking, could feed ahydraulic fire.
328

Use only minimum possible toe brakepressure to hold aircraft stationary
until engine is shut down and nose
wheel is chocked.
3.

EPU switch - OFF.

4.

Throttle - OFF.

T.O. GR1F16CJ1
5.

Nose wheel - Chocked.

6.

MAIN PWR switch - OFF.

7.

Exit toward the front of the aircraft.

If main or standby generator cannot be
reset, NWS is inoperative unless the
EPU is activated.

If a brake fire occurs:
8.

Go to GROUND EGRESS, this section.

6.

Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.

MAIN GENERATOR FAILURE (GROUND)EMERGENCY ENTRANCE AND CREW RESCUE

If the main generator fails on the ground, the standby
generator provides power for full normal braking(both channels) and NWS. Abort the aircraft. Taxiing
is permissible.

Refer to figure 34 for emergency entrance and crew
rescue procedures.

MAIN AND STANDBY GENERATOR FAILURE(GROUND)
If the main and standby generators fail on the
ground, the FLCS PMG and aircraft battery provide
power for full normal braking (both channels). The
EPU should activate and provide power for NWS.
Stop and engage the parking brake prior toattempting to reset the generators.
If main or standby generator resets and further
taxiing is required, brakes should be checked
carefully. Allow the aircraft to begin rolling slowly
and check for normal braking. If normal braking is
inoperative, immediately engage the parking brake.
If MAIN GEN and STBY GEN lights illuminate:
1.

Stop the aircraft.
Turn EPU on, if required, to obtain NWS.

2.

ANTISKID switch-PARKING BRAKE.

3.

OXYGEN-100%.

4.

EPU switch-OFF.

If chocks are not installed, be preparedto immediately engage the parking
brake if it disengages when the EPU isshut off.

EMERGENCY GROUND JETTISON
Ground jettison of the 300gallon, 370gallon, or
600gallon fuel tank(s) results in the tank(s) striking
the ground before the pylon aft pivots release. Thetank(s) will probably rotate horizontally and may
strike the LG. Use EMER STORES JETTISON on the
ground only as a last resort. Refer to EMERGENCYJETTISON, this section.
GROUND EGRESSThe order of accomplishment of ground egress steps
depends on the nature of the emergency. For quickest
ground egress (without jettisoning the canopy), place
the canopy switch up and then prepare for exit while
the canopy is opening. However, if fire or danger of
explosion exists, accomplish steps necessary foregress prior to opening canopy to retain maximum
protection until ready for exit. Disconnect parachuterisers, lapbelt, survival kit, and gsuit. Oxygen andcommunication leads are quickdisconnect. If
required, the canopy can be jettisoned even after it
has been partially or fully opened. If the canopy is
restrained by debris or jammed by crash damage,attempted jettison may result in a portion of the
canopy rocket exhaust entering the cockpit. Thisexhaust may present a heat and blast hazard in the
cockpit; toxic gases are present and 100 percent
oxygen should be used.
1.

Throttle-OFF.

2.

Ejection safety lever-Safe (up).

3.

Harness and personal equipment-Release.

4.

EPU switch-OFF (time permitting).

If further taxiing is required:
5.

ELEC CAUTION RESET button-Depress.
Toe brakes and parking brake are available
with or without the EPU as long as the MAINPWR switch is not moved to OFF.

Exit over the left side (conditions
permitting) to avoid EPU exhaustgases.
329

T.O. GR1F16CJ1Emergency Entrance and Crew Rescue(Typical)
1. If time and conditions permit:

CANOPY SWITCH

a. Insert a 1/4-inch drive socket wrench/speed handle
into canopy handle lock access plug and rotate ccw
to remove plug.

Cb. Insert an 8-inch or longer piece of number 25 drillrod (or 1/8-inch rod) into opening and push inboard
to unlock canopy handle.

1.c.

c. Position external CANOPY switch to UP.

CANOPYSWITCH

Positioning the external CANOPY switch to UP
prior to unlocking the canopy will overheat
the canopy actuator motor or pop the circuit
breaker.

Dd. If canopy is still not open, insert 1/4-inch drive
socket wrench/speed handle into the external
canopy handcrank receptacle and rotate cw(approximately C 52 or D 87 revolutions
required to fully open canopy).

2. Open the canopy emergency release door and extend

1.d.

the canopy jettison D-handle to full length of cable(approximately 6 feet). When the cable tightens, pullhandle hard to jettison the canopy.

The canopy jettisons upward and back toward
the vertical tail with great force. Stand to theside and slightly aft of canopy to full lengthof the cable to avoid canopy rocket blast.

2.

1.a.

ACCESS PLUG

1.b.
DRILL ROD

(Unlocked)(Locked)C

DF CANOPY HANDLE

RIGHT OR LEFT
SIDE OF AIRCRAFT
1F-16X-1-0027X

Figure 34.(Sheet 1)330

T.O. GR1F16CJ1Emergency Entrance and Crew Rescue(Typical)
ALTERNATE
CONFIGURATION(FIXED PITOTS)

FLIP-UP PITOTS(STOWED)SEAWARSINSTALLEDSEAWARSNOT INSTALLED

3. Rotate ejection safety lever located on

EJECTION SAFETY
LEVER (SAFE)

left of seat to full up (vertical) position.

To prevent possible seat ejection
during rescue, rotate ejection safety
lever located on left of seat to fullup (vertical) position.

4. Disconnect crewmember from lapbelt,

3.

g-suit hose, survival kit straps, andparachute risers.

GR1F-16CJ-1-0117X37

Figure 34.(Sheet 2)

331

T.O. GR1F16CJ1
5.

Canopy-Open.

F D Consider canopy jettison so rearseat occupant can egress more rapidly.

FOpening

the canopy with the
MANUAL CANOPY CONTROL hand
crank is extremely difficult. If immediate egress is required, the canopy
should be jettisoned rather thanopened with the handcrank.

Treat any leak as a hydrazine leakuntil investigation proves otherwise.
1.

OXYGEN-100%.

When on the ground:2.

AIR SOURCE knob-OFF (if required).
Consider turning the ECS off to prevent thepossibility of hydrazine fumes or EPU
exhaust gases entering the cockpit.

If canopy does not raise:
6.

OXYGEN-100%.
FIf AIR SOURCE knob is placed to OFF,also turn off nonessential avionicequipment as electronic equipment
may be damaged.

FIf jettison is unsuccessful, heat, blast,F PX III If AIR SOURCE knob is placed toand toxic gas from the rockets may
enter the cockpit.

OFF, OBOGS caution light will illumi
nate. If OXY LOW warning lightilluminates before ground crew ar
rives with oxygen bottle, activate EOS.

FTo prevent the flow of oxygen into the

cockpit after the oxygen hose is
disconnected, do not select EMER.

7.

Canopy-Jettison.

Pulling the CANOPY JETTISON T
handle other than straight out may
cause the handle to jam.
HOT REFUELING EMERGENCY
In the event of a fire or fuel leak/spill while refuelingin hot pit area, refer to FIRE/OVERHEAT/FUELLEAK (GROUND), this section. In the event of fire in
the area of refueling operation (other than in the hot
pit area), have the refueling operation discontinued
and taxi clear.

3.

Taxi to designated isolated parking area (ifrequired) and park aircraft with left wing intowind if possible.

4.

Insure all nonessential personnel are clear.

5.

EPU switch-OFF.

6.

Shut down the engine (after left main wheel ischocked).
NOTE
To prevent sitting in a sealed cockpit(hot) without ECS, consider waiting
for ground crew to arrive with ladderand oxygen bottle prior to shutting
down the engine.

NWS FAILURE/HARDOVER
ACTIVATED EPU/HYDRAZINE LEAK
If landing with an activated EPU or a hydrazine leakis detected while the engine is running:
Inform landing base of hydrazine leak or EPU
operation and request bioenvironmental services
support.

332

NWS failure may be detected by the NWS FAIL
caution light or uncommanded NWS inputs with nocaution light. If NWS FAIL caution light is on, do not
engage NWS. If the NLG strut is overextended, theNWS cannot engage. If the NLG strut overextends
after NWS engagement, NWS becomes disengaged
and the AR/NWS light goes off.

T.O. GR1F16CJ1S Aircraft factors:GW, stores, nature of theemergency, speed at decision point, and importanceof becoming airborne.

NWS malfunctions at any speed may
cause an abrupt turn, tire skidding or
blowout, aircraft tipping, and/or
departure from the prepared surface.
1.

NWS-Disengage.

2.

AR/NWS light-Verify off.

3.

Rudder and brakes-As required.

TAKEOFF EMERGENCIESDELAYED ROTATION
Several factors can cause the airspeed at whichrotation occurs to be greater than that determinedfrom T.O. GR1F16CJ11. As the weight of external
stores carried increases, more nose down moment
must be overcome to rotate for takeoff. Anotherfactor is the application of roll stick force in addition
to aft stick force. Applying a roll input reduces the
maximum trailing edge up position for one
horizontal tail and increased airspeed may berequired to compensate. The last and mostsignificant factor is improper servicing of the nose
gear strut. Improper servicing may not be detectable
during preflight inspection and may cause rotation
speed to increase by up to 15 knots. All of thesefactors combined may add up to 25 knots to thecomputed airspeed for rotation. If pretakeoff flightcontrol checks were normal and the engine isoperating normally (acceleration check normal), theaircraft will rotate above computed rotation speed.
Therefore, takeoff should not be aborted due todelayed rotation until at least takeoff speed is
attained. Notify maintenance after flight if asignificantly delayed rotation occurred.
ABORT
The decision to abort or continue takeoff depends on
many factors. Considerations should include, but not
be limited to, the following:

S Runway factors:Runway remaining, surface

condition (wet, dry, etc.), type and/or number ofbarriers/cables available, obstructions alongside orat the departure end, wind direction and velocity,
and weather and visibility.

S Stopping factors:Maximum antiskid braking,speedbrakes, aerodynamic braking, hook, and dragchute.

Aborting takeoff at high speed with ablown tire may be more dangerous
than continuing takeoff. For heavy
GW takeoffs, an abort at high speedwith a blown tire is extremely dangerous because braking and directional
control are impaired.

FAt high speed (prior to WOW), forward

stick pressure in excess of approxi
mately 2 pounds results in full trailingedge down deflection of the horizontal
tails. This causes excessive loads on
the NLG which can lead to nose tire
failure and possible structural failureof the NLG.

FFailure to use full antiskid braking or

applying brakes with engine above idle
thrust significantly increases thewheel brake temperature and probability of a wheel brake fire.

Normally, with the short takeoff distances of theaircraft, abort is not a problem unless directional
control is a factor (e.g., blown tire). An early decision toabort provides the most favorable circumstances. If
there is any doubt about the ability to stop on the
runway, lower the hook.
Consider aborting after becoming airborne only whensufficient runway is available and flight to a key
position is not possible.

333

T.O. GR1F16CJ1
Aborts above 100 KCAS require diligent adherence tothe procedures in this section for the abort to be
successful. If aborting after rotation, retard throttle to
IDLE and maintain twopoint attitude while applying
maximum wheel braking (maximum pedal pressure(antiskid on) consistent with maintaining directional
control). When wheel brakes become effective, the nose
automatically lowers. After the nosewheel is on the
runway, use maximum effort braking (full aft stick, fullopen speedbrakes, and maximum wheel braking). If
aborting before rotation, retard throttle to IDLE,
maintain threepoint attitude and apply maximum
effort braking if stopping distance is critical. NWS
should be engaged if directional control is a problem.

The hook may miss the cable if theaircraft is not slow enough to compress
the MLG struts sufficiently to makeWOW or if forward stick pressure is
held.
If on fire:5.

NOTE
With engine shut down, NWS is lostand EPU does not activate automati
cally. After hydraulic pressure drops,
braking is available using the brake/
JFS accumulators only. Stop straightahead and engage parking brake.

Consider following hot brake procedures after any
abort. Taxiing after an abort will further increase braketemperature.

FWhen braking absorbs a high amountof energy, do not shut down engine
until firefighting equipment is avail
able and do not use the parking brake.

FHot wheels and brakes may ignite

leaking hydraulic fluid. Wheel fusible
plugs may relieve tire pressure within
15 minutes after stop.

1.

Throttle-IDLE.

When the throttle is retarded to IDLE
from MAX AB, the thrust and rpm
decay to idle can take up to 24 seconds.
Do not mistake high thrust/rpm for
failure of the engine to respond to theidle command. Engine shutdown fromMAX AB may result in a tailpipe fire.

Throttle-OFF.

6.

FUEL MASTER switch-OFF.

ENGINE MALFUNCTION ON TAKEOFF
An engine malfunction on takeoff presents ademanding situation where critical actions must be
accomplished quickly with little time for analysis. If
takeoff is continued, a straight ahead climb isgenerally preferred over an immediate turn to low
key. This action provides more favorable ejection
parameters and an increase in analysis time. If
necessary, use only shallow turns to avoid
aggravating the situation. Jettison stores if required
to reduce GW.
ENGINE FAILURE ON TAKEOFF

2.

DRAG CHUTE switch-DEPLOY (if required).

Engine failure shortly after liftoff may not permit
time for analysis or corrective action. The primaryconcern should be to trade any excess airspeed for
altitude and to eject prior to allowing a sink rate todevelop. Jettisoning stores may aid in gaining
altitude but must not delay the ejection decision. If
the failure occurs later in the takeoff phase, time may
be available for analysis or corrective action.

3.

Wheel brakes-Apply (as required).

If conditions permit:

4.

HOOK switch-DN (if required).

1.

S The hook should be lowered at least 1500 feet

If conditions do not permit an abort:from the cable to allow adequate time for hook
to stabilize and for full holddown force to be
developed by the hook actuator.

S Refer to CABLE ARRESTMENT, this section.

334

Abort.

1.

Zoom.

2.

Stores-Jettison (if possible).

3.

Eject.

T.O. GR1F16CJ1
AB MALFUNCTION ON TAKEOFF

If takeoff is continued and/or thrust is insufficient:

An AB malfunction can be detected by a thrust loss
and nozzle closure or failure of AB to light withinallowed time or stalls accompanied by a loud bangor pop. An AB failure (other than a slow/no light)
may indicate other engine problems. If possible,
abort the takeoff. If takeoff is continued, the throttle
should be retarded to MIL. If normal thrust is notavailable in MIL, refer to LOW THRUST ON
TAKEOFF OR AT LOW ALTITUDE (NONAB), thissection. AB operation should not be reattempted
unless required to sustain flight.

1.

If thrust is still insufficient or AB does not light:
2.

Abort.

If takeoff is continued:
1.

Throttle-MIL.

2.

Stores-Jettison (if required).

LOW THRUST ON TAKEOFF OR AT LOW ALTITUDE(NONAB) PW 229
Low altitude, for engine malfunction purposes, isgenerally defined as 10,000 feet AGL or below.
Low thrust can be the result of DEECrelated
failures; a failed open, damaged or missing nozzle; oran engine rpm rollback. A failed open, damaged or
missing nozzle may result in significant thrust loss
and the inability to take off or maintain level flight.
For description of failed open, damaged or missing
nozzle, refer to NOZZLE FAILURE PW 229 , thissection. Low thrust can also be the result of the start
bleed strap failing to close during the normal start
cycle.
If low thrust occurs during takeoff and conditions
permit, the takeoff should be aborted. If the takeoff
must be continued or in any critical phase of flightand MIL thrust is not sufficient, AB should be used.
An excessively open nozzle may reduce the chancefor successful AB light. If the AB does not light(allow the DEEC to automatically resequence theAB if conditions permit), place the ENG CONT
switch to SEC.
If an automatic transfer to SEC occurs or SEC isselected manually, resulting thrust is 7080 percent
of normal MIL thrust with no AB capability. If
thrust is still low, consider jettisoning stores.
If on takeoff and the decision is made to stop:
1.

Abort.

ENG CONT switch-SEC.

With nozzle loss, catastrophic engine
failure and fire are probable withprolonged high power settings above
850_C FTIT while in SEC.

If decision is made to stop:
1.

Throttle-AB.
The chances for a successful AB light with thenozzle open more than 30 percent are
reduced.

NOTE
In a partial thrust situation, thrustavailable may increase as altitudedecreases. 250 knots approximates theairspeed at which thrust required forlevel flight is the lowest.
3.

Stores-Jettison (if required).

If nozzle is failed open, damaged, or missing:4.

Airspeed-Climb to arrive at 250 knots ordescend at 250 knots to obtain level flight
above minimum recommended ejection altitude
or minimum safe altitude, whichever isappropriate.
NOTE
FWith a missing nozzle, level flight may

not be attainable above 5000 feet MSL.

FIf descent is required, maintain 250

knots with throttle set at 850_C FTIT.

If level flight cannot be maintained by 1000 feet above
minimum recommended ejection altitude or minimum safe altitude, whichever is appropriate:
5.

Throttle-As required to maintain 250 knots
in level flight.

If airspeed drops below 250 knots,
trade altitude to reacquire 250 knots.
Do not descend below minimum
recommended ejection altitude or
minimum safe altitude, whichever isappropriate.
335

T.O. GR1F16CJ1
6.

Land as soon as possible. Plan a flameout
landing. Refer to FLAMEOUT LANDING, thissection.

LOW THRUST ON TAKEOFF OR AT LOW ALTITUDE(NONAB) GE 129
Low altitude, for engine malfunction purposes, isgenerally defined as 10,000 feet AGL or below.
Low thrust can be the result of DECrelated failures,
the nozzle failing, or an rpm rollback. Thesesituations may result in significant thrust loss andthe inability to take off or maintain level flight. If low
thrust occurs during takeoff and conditions permit,
the takeoff should be aborted.

ENGINE FIRE ON TAKEOFF
An engine fire may be indicated by the ENG FIRE
warning and/or OVERHEAT caution lights, high
FTIT, smoke, or fumes. Refer to ENGINE FIRE, thissection.
LG FAILS TO RETRACT
If the LG handle warning light remains on after the
LG handle is placed to UP, the LG or LG doors are notfully up and locked.
1.

Airspeed-300 knots maximum.

2.

LG handle-DN. (Use DN LOCK REL button
if required.)If the takeoff must be continued or in any criticalphase of flight, when MIL thrust is insufficient, AB
should be used. An automatic transfer to HYB or SEC
may occur resulting in less than MIL thrust with noAB capability (SEC). An excessively open nozzle may
reduce the chance for a successful AB light. If the AB
does not light, the ENG CONT switch should beplaced to SEC.
If an automatic transfer to SEC occurs or SEC isselected manually, resulting thrust is 7095percent of
normal MIL thrust with no AB capability. If thrust is
still low, consider jettisoning stores.
If on takeoff and the decision is made to stop:
1.

If LG handle does not lower, selectBRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND. PW 229
Nozzle remains closed, resulting inhigher than normal landing thrust.
If LG comes down normally:
3.

GW-Reduce prior to landing.

If LG does not indicate down:

Abort.

If takeoff is continued and/or thrust is insufficient:
1.

Throttle-AB.
The chances for a successful AB light with thenozzle open more than 30 percent are
reduced.

If thrust is still insufficient or AB does not light:
2.

ENG CONT switch-SEC.
NOTE
In a partial thrust situation, thrustavailable may increase as altitudedecreases. 250 knots approximates theairspeed at which thrust required forlevel flight is the lowest.

3.

Stores-Jettison (if required).

336

Do not cycle LG handle. Damage to LG
or LG doors may result.
4.

Go to ALTERNATE LG EXTENSION, thissection.

LG HANDLE WILL NOT RAISE
If the left MLG WOW switch fails to the groundposition, the LG handle does not move out of the DN
position. In addition, the TO/LDG CONFIG warning
light and touchdown skid control system are affected.
The LG handle may be raised by first depressing the
LG handle downlock release button.
If conditions permit:
1.

Airspeed-300 knots maximum.

2.

GW-Reduce prior to landing.

T.O. GR1F16CJ1
If LG must be raised:1.

LG handle DN LOCK REL button-Depress.

2.

LG handle-UP.
TO/LDG CONFIG light is on if left MLG
WOW switch has failed.

NOTE
The decision to take off or abort depends
on the speed at the time of the failure,
GW, stopping distance required, andarresting gear availability.

When desired:3.

LG handle-DN. (Use DN LOCK REL button
if required.)

If LG handle does not lower, selectBRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND. PW 229
Nozzle remains closed, resulting inhigher than normal landing thrust.
After touchdown:
4.

Brakes-Apply after wheels spin up.

Touchdown antiskid protection may
not be available. Landing with feet on
the brake pedals may result in blowntire(s).

If takeoff is continued, do not retract the LG, reduce
GW if practical, and prepare to land as soon aspractical.
Directional control during stopping is the primaryconcern when aborting with a blown tire. Heavy GW
and high speed aborts place greater demands on thebrakes and tires. This may cause damage to the NWS,
wheels, and struts which may result in loss of
directional control. In addition, heavy differential
braking may result in MLG tire failure.
If aborting with a blown MLG tire, leave antiskid on
to minimize possibility of skidding the good tire. If thewheel with the blown tire does not turn, the antiskid
switches to the b2t alternate braking mode, LESS b2t
pulsating antiskid mode. Use roll control to relievepressure on the blown tire and NWS to maintaindirectional control.

Possible indications of a NLG tire failure include aloud explosion, slight deceleration, vibrations, flying
debris, and at night, a flash or flame. These
characteristics can be mistaken for an engine stall.
Rubber debris may cause damage to the engine, NWS
wiring harness, WOW switch assembly and/or gear
position sensor wiring. NWS may not be available
even though the AR/NWS light is on and the NWS
FAIL light is off.

If aborting with a blown NLG tire, hold the nosewheel
off the runway (if able) and use twopoint
aerodynamic braking until control effectivenessbegins to decay. Lower the nosewheel to the runway
and immediately engage NWS, if available, tomaintain directional control. Use aft stick to reduceload on the NLG after brakes are applied. If NWS is
not available, the aircraft tends to drift right.
Attempt to move to the left side of the runway before
rudder effectiveness is lost and maintain directional
control with rudder and differential braking. Stop
short of the departureend arresting cable if possible.
The small nosewheel rolling radius with the tire
missing may allow the cable to pass over top of thenosewheel and cause NLG collapse.

Aborting takeoff at high speed with ablown tire may be more dangerous
than continuing takeoff. For heavy
weight takeoffs, an abort at high speedwith a blown tire is extremely dangerous because braking and directional
control are impaired.

A NLG tire failure accompanied by complete tire
separation from the wheel may cause reverse
castering. The conditions for this to occur are theNWS disengaged or inoperative; the nose wheel rim
rolling on a deformable surface (i.e. asphalt); andlateral force applied to the nose wheel from either arudder input or differential braking. If reverse
castering occurs the nose wheel will turn in the
opposite direction of rudder and brake inputs making
it extremely difficult to maintain directional control.

BLOWN TIRE ON TAKEOFF
Tire failure on takeoff is difficult to recognize and may
not be noticed in the cockpit.

Change 1337

T.O. GR1F16CJ1CANOPY WARNING LIGHT ON
If CANOPY warning light illuminates:
If a blown NLG tire occurred and NWS
is not available, it may not be possible toprevent departure from the runway. A
reverse castering effect may occur inwhich the nosewheel moves opposite to
the rudder or differential braking input.

1.

Canopy handle - Push outboard.

If CANOPY warning light remains on:2.

Go to CANOPY LOSS/PENETRATION IN
FLIGHT, this section.

CANOPY LOSS/PENETRATION IN FLIGHT

With a blown tire, avoid centerline
lights as they may cause wheel damage
and subsequent loss of directional
control. Failure to use full aft stickwith a blown NLG tire may lead towheel failure and directional control
problems.

Canopy loss/penetration in flight results indisorientation and may result in structural damage
caused by the canopy striking the aircraft. Due tothe possibility of severe disorientation, vision loss,
injury, or incapacitation at high airspeed, immediate ejection may be the only option. Slow to 180
knots or less and check for controllability. Wind
blasts up to 180 knots can be coped with by leaning
forward and down behind the glareshield and HUD.

Stop straight ahead and shut down the engine as soon
as firefighting equipment is available. Do not attempt
to taxi unless an emergency situation exists.

FArms must be kept close to the body to

If takeoff is not feasible:
1.

avoid letting wind blast pull arms outof the cockpit.

Abort.

FHUD

glass disintegration can be
expected following a medium to highenergy bird strike with or without
canopy penetration.

If takeoff is continued:
1.

LG - Do not retract.

2.

Airspeed - 300 knots maximum.

3.

Refer to LANDING WITH A BLOWN TIRE,
this section.

INFLIGHT EMERGENCIES
When preparing to activate backup systems which
rely on stored nitrogen pressure to function, consider
the potential for timerelated failures and do not
activate the system earlier than required. For
example, the hydrazine mode of the EPU requires
nitrogen pressure to force hydrazine to the EPU. If anitrogen leak exists, turning the EPU on early could
lead to an inability of the EPU to function on
hydrazine. Similarly, alternate LG and hook extension
use stored nitrogen pressure. If alternate LG
extension is used early and a nitrogen leak exists,
hydraulic system B could subsequently fail. Such aleak could also result in insufficient pressure to
maintain proper hook holddown force. Since nitrogenleaks are not apparent prior to system activation,
consider their potential existence and activate thebackup system when needed, but not excessively early.
338Change 1

Wind buffet increases slightly with increased AOA.
Therefore, if fuel is not critical, TEF's should be
extended using the ALT FLAPS switch or by placing
the LG handle to DN.
1.

Airspeed-180 knots maximum.

2.

Seat-Full down.

3.

ALT FLAPS switch-EXTEND.

4.

Land as soon as possible.

DRAG CHUTE DEPLOYED IN FLIGHT
If the drag chute is deployed in flight below 190 knots:
NOTE
If the drag chute is deployed below
approximately 190 knots, it does notbreak away from the aircraft.
1.

DRAG CHUTE switch - REL.

If the drag chute does not release:2.

Throttle - MAX AB.

T.O. GR1F16CJ1
COCKPIT PRESSURE/TEMPERATURE MALFUNC
TION
Loss of cockpit pressurization could be caused bycanopy seal, airconditioning system, or cockpit
pressure regulator safety valve malfunctions or
ECS shutdown or failure.
Certain ECS equipment malfunctions result in
temporary shutdown of the ECS. These shutdowns
are more prevalent at high altitude during low
speed flight with high engine thrust settings. An
ECS shutdown is characterized by an oily, smokey
smell, followed by loss of cockpit noise and airflow
and gradual loss of pressurization. These tempo
rary shutdowns typically last from 2045 seconds or,on occasion, up to 2 minutes. The EQUIP HOT
caution light may illuminate if the shutdown lasts
longer than 20 seconds.

If the cockpit temperature is excessive and does not
respond to AUTO or MAN temperature commands
or cockpit pressure is lost, proceed as follows:
1.

OXYGEN-100%

2.

Altitude-25,000 feet maximum.

3.

Airspeed-500 knots maximum.

4.

AIR SOURCE knob-OFF (1015 seconds),
then NORM.
PX III The OBOGS caution light illuminates
while AIR SOURCE knob is in OFF.

If cockpit pressure is not regained but all othersystems dependent on the ECS are operational:
5

Flight may be continued below 25,000 feet.

If ECS has failed or cockpit temperature control is notregained:
FWith the ECS shut down or the AIR5

AIR SOURCE knob-OFF.

SOURCE knob in OFF or RAM, thegsuit does not inflate and PBG isdisabled.

F PX III With the ECS shut down or the

PX III If AIR SOURCE knob is placed to
OFF or RAM, OBOGS is inoperative.
Activate EOS if OXY LOW warning
light illuminates above 10,000 feetcockpit altitude.

AIR SOURCE knob in OFF or RAM,OBOGS is inoperative. Activate EOS if
OXY LOW warning light illuminates
above 10,000 feet cockpit altitude.

Short duration (approximately 15 seconds) losses of
cockpit airflow when operating above 35,000 feet
MSL should not be confused with an ECS shutdown.
These are the result of an automatic ECS cutback
which is designed to prevent total system
shutdown.
NOTE
PX III The OBOGS caution light may
illuminate as a result of ECS cycling or
temporary ECS shutdown. This isnormal as long as the OXY LOW
warning light does not illuminate.

Most AUTO position temperature failures can becorrected by use of the MAN position.
If cockpit pressure altitude exceeds 27,000 feet, the
CABIN PRESS caution light illuminates.

6.

AIR SOURCE knob-RAM (after cockpit isdepressurized).

PX III If AIR SOURCE knob is placed to
OFF or RAM, OBOGS is inoperative.
Activate EOS if OXY LOW warning
light illuminates above 10,000 feetcockpit altitude.

NOTE
External fuel cannot be transferred in
OFF or RAM. Consider jettisoning
tank(s) to decrease drag if range iscritical and ECS cannot be turned on
for short periods of time to transferfuel.

339

T.O. GR1F16CJ1
7.

Nonessential electrical equipment-Off.

NOTE

NOTE

If in VMC and the ADI and HSI are notrequired for flight, the EGI/INS should
be considered nonessential.

If in VMC and the ADI and HSI are notrequired for flight, the EGI/INS should
be considered nonessential.

4.

8.

Land as soon as practical.

EJECTION

9.

Check for failed emergency dc bus(es). Refer to
EMERGENCY POWER DISTRIBUTION, thissection.

Refer to figure 34.1. Ejection should be accomplishedat the lowest practical airspeed.

Land as soon as practical.

EQUIP HOT CAUTION LIGHT
If EQUIP HOT caution light illuminates:
NOTE
FCertain ECS equipment malfunctions

result in temporary shutdown of the
ECS and illumination of the EQUIP
HOT caution light.

FAn ECS shutdown and EQUIP HOT

caution light illumination for up to 2
minutes can occur either duringextended LG down flight between sea
level and 7000 feet MSL or during
operation above a line from 42,000 feet
MSL at 0.2 mach to 50,000 feet MSL at
0.95 mach. These ECS shutdowns arenormal, but may still require addition
al action if the EQUIP HOT lightremains on for more than 1 minute.

FIf cockpit temperature is excessive,refer to COCKPIT PRESSURE/TEM
PERATURE MALFUNCTION, thissection.

1.

AIR SOURCE knob-Confirm in NORM if
smoke or fumes are not present.

2.

Throttle-80 percent rpm minimum (inflight).

If EQUIP HOT caution light remains on after 1
minute:3.

Nonessential avionics-Off.

340Change 1

FThe minimum altitude obtained fromEJECTION SEAT PERFORMANCE
charts, Section I, does not include any
allowance for pilot decision making,changing flight conditions, or hand
movement from the stick and/or
throttle to the ejection handle. There
fore, minimum altitude for ejection
decision could be significantly higher.

FWhen in a spin/deep stall or otheruncontrolled flight, eject at least 6000
feet AGL whenever possible. This is
the minimum altitude to initiate
ejection with minimal risk of injury
under the most adverse conditions.
The decision to eject must have been
made prior to this altitude. Delaying
ejection below this altitude may result
in serious injury or death.

FUnder controlled flight conditions,eject at least 2000 feet AGL wheneverpossible. If below 2000 feet AGL,
attempt to gain altitude if airspeed
permits. Do not delay ejection below
2000 feet AGL for any reason which
may commit you to unsafe ejection.

FFailure to monitor sink rate and height

above terrain while performing anairstart or applying low thrust recov
ery procedures can result in an ejection
outside the ejection seat performanceenvelope.

T.O. GR1F16CJ1

F-16 ACES II Ejection Injury Risk Chart
245
240
INCREASED RISKLIMB FLAIL
400 KTS

220

INCREASED RISKSEAT STRUCTURE
FAILURE
500 KTSBODY WEIGHT (LBS)

EJECTION SEAT DESIGN LIMITS

200

180

160

HIGH RISK.
AVG.
MAJOR/FATALINJURY RISK:
80%

MODERATE RISK
AVG. MAJOR/FATALINJURY RISK: 36%

LOW RISK
AVG. MAJOR INJURY
RISK: 9%EJECTION SEAT DESIGN LIMITS

140

INCREASED RISKLIMB FLAIL AND DROGUEOPENING SHOCK
420 KTS

120
103
0

100

200

300

400

AIRSPEED (KTS)

500

600

700
1F-16X-1-4023X

Figure 34.1

Change 1340.1/(340.2 blank)

T.O. GR1F16CJ1Ejection (Immediate)
1.
FThe ACES II ejection seat was designed

for body weights in the 140 to 211 pound
range. There are additional ejection
injury risks associated with bodyweights outside this range.
DForbody weights less than 140
pounds, limb flailing, less seat stabil
ity, and more severe drogue chuteopening shock (ejection modes 2 and
3) are concerns. The risk of injury
associated with limb flailing anddrogue chute opening shock increases
for ejection above 420 knots. This
injury risk also increases as bodyweight decreases below 140 pounds.

Ejection handle-Pull.

Ejection (Time Permitting)If time permits, descend to avoid the hazards of highaltitude ejection. Stow all loose equipment and direct
the aircraft away from populated areas. Sit with headagainst headrest, buttocks against back of seat, andfeet on rudder pedals.
1.

IFF MASTER knob-EMER.

2.

Loose equipment and checklist-Stow.

3.

Lapbelt and helmet chin strap-Tighten.

4.

Night vision devices-Remove (if appropriate).

D For body weights greater than 211

pounds, limb flailing, seat structural
failure, and parachute landings areconcerns. The risk of injury from limb
flailing is high for ejection above 400
knots. The seat leg braces frequently
deform during ejections above 500
knots; this deformation has led toseat side panel failures (and unsuc
cessful ejections) during 600 knot
ejection tests. The risk of injury
during parachute landing is threetimes the average. These injury risks
also increase as body weight in
creases above 211 pounds.

FWind blast exerts medium force on the

body up to 400 knots, severe forces
causing flailing and skin injuries
between 400600 knots, and excessive
force above 600 knots.

FDuring high altitude ejections (modeFailure to remove night vision goggles(NVG) prior to ejection may cause
serious injury. If unable to remove NVG,
a proper ejection body position (headback against the seat headrest) reduces
the chance of injury from the NVG.
5.

Visor-Down.

6.

Throttle-IDLE.
Slow to lowest practical airspeed.

7.

Assume ejection position.

8.

Ejection handle-Pull.

Failure of Canopy To Separate
If canopy fails to separate, remain in position forejection while keeping arms inboard and perform thefollowing:3), automatic pilot/seat separation andrecovery parachute deployment occur
between 16,00014,500 feet MSL. Ifhigh terrain is a factor, manual seat
separation procedures must be used to
bypass the automatic sequence.

To eject, grasp ejection handle using a twohanded
grip with thumb and at least two fingers of each hand.
Pull up on handle and continue holding until
pilot/seat separation. The ejection handle does notseparate from the seat.
Refer to figure 35 for manual seat separation andmanual survival equipment deployment.

If canopy is jettisoned or manually
released/opened after pulling theejection handle, the ejection seatfunctions immediately after canopy
separation. Be prepared to immediate
ly put arm back in ejection position
when the canopy starts to separate.
1.

Canopy - Open normally.

2.

Canopy - Jettison.
Change 1341

T.O. GR1F16CJ1Manual Survival Equipment Deployment/
Manual Seat Separation
SURVIVAL EQUIPMENT (TYPICAL)
A. IF EMERGENCY OXYGEN FAILS TO ACTIVATE AUTOMATICALLYUPON EJECTION, PULL THE EMERGENCY OXYGEN GREEN
RING LOCATED NEAR THE LEFT HIP.

B. AFTER RECOVERY PARACHUTE ASSEMBLY DEPLOYMENT, RAISEC.

VISOR AND DISCARD OXYGEN MASK. IF SURVIVAL KIT DOESNOT DEPLOY AUTOMATICALLY, GRASP KIT RIPCORD HANDLE
WITH RIGHT HAND AND PULL. KIT RIPCORD HANDLE IS LOCATEDNEAR RIGHT HIP.

C. LIFERAFT INFLATION IS INITIATED WHEN THE DROP LINE/LANYARD IS FULLY EXTENDED AFTER SURVIVAL KITDEPLOYMENT. CHECK LIFERAFT AND IF NOT INFLATED,SNATCH PULL DROP LINE/LANYARD TO INFLATE.

NOTE
If the survival kit is deployed after landing inwater, a snatch pull on the drop line/lanyard (nearCO 2 bottle) is required to inflate the liferaft.

A.
B.

MANUAL SEAT SEPARATIONTO PERFORM MANUAL SEAT SEPARATION AND DEPLOY THE RECOVERY PARACHUTE ASSEMBLY, PULL THE EMERGENCY MANUAL CHUTE HANDLE.

After ejection, the EMERGENCY MANUAL CHUTE
handle should only be used if the automaticsequence has failed or if high terrain is a factor.
Pilot/seat separation in modes 1 and 2 should
occur rapidly after pulling the ejection handle.
If the pilot has time to realize seat separation
has not taken place, a failure has probably
occurred and manual seat separation should be
performed. In mode 3, pilot/seat separation
occurs between 16,000-14,500 feet MSL. If
automatic pilot/seat separation does not occur
in this altitude range or if high terrain is a factor,
manual seat separation must be performed.
Failure to fully pull the EMERGENCY MANUALCHUTE handle may result in ballistically deploying
the recovery parachute assembly without releasing
the lapbelt and inertia reel straps and unlatching
the seat pan lid.
Do not attempt to open the lapbelt. If the lapbelt is opened, the seat will partially fall away,
but the parachute risers remain attached to theinertia reel straps. The only way to separate fromthe seat is to pull the EMERGENCY MANUAL CHUTE
handle at least 6 inches.

Figure 35.

342

1F-16X-1-2023X

T.O. GR1F16CJ1Single Generator Failures (in Flight)

Pulling the CANOPY JETTISON
Thandle other than straight out may
cause the handle to jam.
3.

MANUAL CANOPY CONTROL handcrank Push in and rotate ccw.

Use of the CANOPY JETTISON T
handle or MANUAL CANOPY CON
TROL handcrank may result in seri
ous injury. To minimize chances of
injury, immediately release the handle
when the canopy starts to separate.
Ejection Seat Failure
If the ejection seat fails to function after the ejection
handle is pulled and the canopy has separated fromthe aircraft, there are no provisions designed into
the escape system for manual bailout.
DITCHING
Ditch the aircraft only as a last resort. All attempts
to eject should be accomplished prior to ditching.
ELECTRICAL SYSTEM FAILURES
Electrical system failures are indicated by illumina
tion of the ELEC SYS caution light and one or more
ELEC control panel lights (in any combination). Afteraccomplishing the appropriate emergency procedures, refer to EMERGENCY POWER DISTRIBU
TION, this section, to determine inoperativeequipment for any remaining ELEC control panel
lights.
FLCS PMG Failure
If all four branches of the FLCS PMG fail in flight,
the FLCS PMG light illuminates. The converter/regulator automatically selects the power source withthe highest voltage from the available alternate
sources. Other FLCS power sources are the main
generator, the standby generator, the EPU generator,
the EPU PMG, and the aircraft battery.
If FLCS PMG light illuminates:
1.

Land as soon as practical.

A single generator failure is indicated by illumina
tion of either the MAIN GEN or STBY GEN light.
A mechanical failure which affects the standby
generator may also affect the FLCS PMG causing
illumination of both the STBY GEN and FLCS
PMG lights. Actions required in this circumstanceare the same as those for illumination of the MAIN
GEN or STBY GEN light; thus, this failure is also
categorized as a single generator failure.
Illumination of the MAIN GEN light indicates that
one or both nonessential ac buses are not being
powered by the main generator. The light may
illuminate if the main generator fails, if a main
power contactor trips off one or both nonessential
ac buses, or if a shorting failure of an overcurrent
sensing contactor (OCSC) or other wiring/equipment
occurs. Refer to PARTIAL ELECTRICAL POWER
LOSS, this section, for discussion of OCSC failures. In
all cases, the standby generator should automaticallycome on line to power the essential and emergency
buses which are not being powered by the main
generator. The nonessential dc bus is not powered.
The main generator may be reset by depressing the
ELEC CAUTION RESET button or cycling the MAINPWR switch; however, moving the switch out of
MAIN PWR removes standby generator power andcauses the EPU to activate. A reset attempt should
not be made if the MAIN GEN light was preceded bya 23 second loss of power to the HUD, MFD's, andother cockpit equipment. Refer to EMERGENCY
POWER DISTRIBUTION, this section, for a list of
inoperative equipment if the main generator is/does
not reset.
Illumination of the STBY GEN light indicates thatstandby generator power is not available for the
essential and emergency buses. The STBY GEN lightilluminates for standby generator system failure orfor failure of the essential bus relay(s). No bus loses
power if the STBY GEN light illuminates when themain generator is still on line. The standby generator
may only be reset by depressing the ELEC CAUTIONRESET button.
Illumination of both STBY GEN and FLCS PMG
lights indicates that standby generator and FLCS
PMG power is not available. If the main generator
is on line, no buses lose power and the EPU does not
automatically activate.

343

T.O. GR1F16CJ1
NOTE

Illumination of the MAIN GEN lightafter a 23 second loss of power to the
HUD, MFD's, and other cockpit instru
ments indicates shorting failure of an
OCSC or other wiring/equipment.
NOTE

PX II The VHF radio is not powered
when the EPU is running or the main
generator is off line.

If MAIN GEN and STBY GEN lights illuminate:
1.

EPU switch - ON (if EPU run light is off).

2.

ELEC CAUTION RESET button - Depress.
This action may reset the main and/or
standby generator. The MAIN PWR switch
may also be cycled to reset the main
generator.

F PX II The VHF radio is not powered

when the main generator is off line.

FWith standby generator failure andthe MAL & IND LTS switch in DIM,
the ELEC SYS caution light may notappear to illuminate when the MASTER CAUTION and STBY GEN lights
illuminate.

If MAIN GEN or STBY GEN light goes off:3.

EPU switch - OFF, then NORM.

4.

ADI - Check for presence of OFF and/or AUX
warning flags.
If warning flag(s) is in view, refer to EGIFAILURE/TOTAL INS FAILURE, this section.

If MAIN GEN light illuminated after a 23 second
loss of the HUD and MFD's was observed:
1.

Land as soon as practical.

If MAIN GEN light illuminated and a 23 second lossof the HUD and MFD's was not observed, or if STBYGEN or STBY GEN and FLCS PMG lights illuminate:
1.

2.

PX II If only AUX flag is in view, pitch
and roll attitude information is likely
to be erroneous due to INS autorestartin the attitude mode when other thanstraight and level, unaccelerated
flight conditions existed.

ELEC CAUTION RESET button - Depress.
This action may reset the main or standby
generator. Cycling the MAIN PWR switch
may also reset the main generator; however,this action momentarily removes standby
generator power and activates the EPU.

5.

Land as soon as practical.

Land as soon as practical.

6.

Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.

Main and Standby Generator Failure (in Flight)
Illumination of the MAIN GEN and STBY GEN
lights indicates that power is lost from at least one
nonessential ac bus and that the standby generator
also failed. The EPU should start automatically at
the time of the second failure as indicated by the
EPU AIR and EPU run lights. Additional lights,
such as HYDRAZN, AVIONICS FAULT, andENGINE FAULT, may also illuminate. If bothgenerators fail, power to all essential and non
essential buses (ac and dc) is lost. Refer to
EMERGENCY POWER DISTRIBUTION, this section, for list of inoperative equipment. The EPU
generator powers the emergency ac and dc buses andboth battery buses. Depressing the ELEC CAUTIONRESET button may reset the main and/or standby
generator.
344Change 1

If MAIN GEN and STBY GEN lights remain on:
3.

ADI - Check for presence of OFF and/or AUX
warning flags.
If warning flag(s) is in view, refer to EGIFAILURE/TOTAL INS FAILURE, this section.

PX II If only AUX flag is in view, pitch
and roll attitude information is likely
to be erroneous due to INS autorestartin the attitude mode when other thanstraight and level, unaccelerated
flight conditions existed.

T.O. GR1F16CJ1
4.

Land as soon as possible.

5.

Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.

6.

If hydrazine depletes or EPU run light goes off
at low thrust - Go to ABNORMAL EPU
OPERATION, this section.

After landing and aircraft is stopped:
7.

Chocks - Installed (or parking brake engaged).

8.

EPU switch - OFF.

If chocks are not installed, be preparedto immediately engage the parking
brake if it disengages when the EPU isshut off.
9.

MAIN PWR switch - MAIN PWR (until chocks
are installed).

Main, Standby, and EPU Generator Failure
A main, standby, and EPU generator failure isindicated by illumination of the MAIN GEN and
STBY GEN lights without an EPU run light. Other
indications include loss of all avionics, ADI OFF and
AUX warning flags, and uncontrollable cold airflow
into the cockpit or reduced airflow to the cockpit if
the water separator coalescer freezes up, and PX III
loss of OBOGS. The caution lights which illuminate
for a failure of just the main and standby generators(e.g., AVIONICS FAULT, ENGINE FAULT) do notilluminate. Refer to EMERGENCY POWER DISTRIBUTION, this section, for a list of inoperativeequipment. The primary concerns during multiple
generator failures are maintaining power to the
FLCS and maintaining/regaining power to theemergency buses.

PX III With a main, standby, and EPU
generator failure, OBOGS and the OXY
LOW warning light are inoperative.
Activate EOS if above 10,000 feetcockpit altitude.

The EPU generator may be inoperative for several
reasons, two of which may be remedied from the
cockpit. If the EPU AIR light is off, the EPU may nothave received an automatic start command;
manually turning the EPU on may correct thisfailure. If the EPU GEN, EPU AIR, and HYDRAZN
lights are illuminated but the EPU run light is off,the EPU may be underspeeding. If the EPU PMG
light is on or blinking, EPU speed is very slow. The
underspeed could be caused by failure of hydrazine
to power the EPU in conjunction with a low thrust
setting and may be corrected by advancing thethrottle. If the EPU generator operates, refer to
MAIN AND STANDBY GENERATOR FAILURE (IN
FLIGHT), this section.
If the EPU generator is still inoperative and themain and standby generators do not reset, power
cannot be restored to the emergency buses. The
aircraft battery supplies power to battery buses No.
1 and No. 2. As long as the ACFT BATT TO FLCS
light remains off, the EPU PMG and/or FLCS PMG
is supplying adequate FLCS power. The primaryconcern is aircraft battery life for communications,brakes, and hook. UHF communications areavailable on the frequency in use at power loss with
UFC selected. BACKUP must be selected if afrequency change is desired.
If the FLCS PMG light illuminates when the MAIN
GEN, STBY GEN, and EPU GEN lights are on, there
are two possible sources of FLCS power (EPU PMG
and aircraft battery). If the EPU PMG and ACFT
BATT TO FLCS lights are off, the EPU PMG is
supplying power to the FLCS whether the EPU run
light is on or off.
If the ACFT BATT TO FLCS light is on, the aircraft
battery is supplying power to the FLCS. The aircraft
battery supplies power for FLCS operation for 314
minutes (depending on state of battery charge) after
total generator failure. As the aircraft battery
depletes, the flight controls either become unrespon
sive or uncommanded maneuvers occur.

Imminent loss of electrical power tothe FLCS is indicated by increasingly
degraded flight control response anduncommanded motions. Total loss of
FLCS power results in a pitchingmotion and complete loss of control. At
high airspeeds, the pitching motion
could be excessive and interfere withthe ability to eject.
345

T.O. GR1F16CJ1
NOTE

3.

Climb if necessary.

If total loss of FLCS power occurs in thelanding configuration and near final
approach airspeed, the pitching motion is gradual and in the noseup
direction for all configurations.

4.

Throttle - As required to extinguish the
HYDRAZN light.

The ACFT BATT FAIL light indicates battery
voltage less than 20 volts. Brake operation is
doubtful during total generator failure with the
ACFT BATT FAIL light on. As aircraft battery
voltage continues to decrease, the capability tooperate the brakes, lower the hook, and deploy thedrag chute is lost. Lower the hook early sincesignificantly higher battery voltage is required tolower the hook than is required to keep it fully
extended. Once lowered, the hook remains full downwell past the point at which the brakes are lost. An
approachend arrestment is recommended, if
conditions permit, because it is difficult to ascertain
brake operation. Relative intensity of the warningand caution lights is not a positive indication ofbattery voltage level.
If MAIN GEN, STBY GEN, and EPU GEN lights
illuminate:

If EPU GEN light goes off:5.

If EPU GEN light is still on:
6.

7.

8.

EPU switch - OFF, then NORM.

9.

Land as soon as possible.

10. Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.
If MAIN GEN, STBY GEN, and EPU GEN lights all
remain on or all come on again:

FPlan to land within 30 minutes toinsure adequate electrical power forcommunications, brakes, hook, anddrag chute.

FIf the FLCS PMG and EPU PMG lights

are on in combination with the ACFT
BATT TO FLCS light, the aircraft battery
is powering the FLCS. With the aircraft
battery powering the FLCS in addition to
the battery buses, approximately 314
minutes flight time is available.

The VHF radio is not powered
when the EPU is running or the main
generator is off line.
PX II

2.

AOA - 12 degrees maximum (200 knots
minimum).

EPU switch - ON (if EPU run light is off).

346

MAIN PWR switch - BATT, then MAIN PWR.
This action may reset the main generator.

If either MAIN GEN or STBY GEN light goes off:

NOTE

LEF's are inoperative and departure
susceptibility may be increased. Near 1g
flight, 200 knots should keep AOA lessthan 12 degrees. Limit rolling maneuvers
to a maximum bank angle change of 90
degrees and avoid rapid roll rates.

ELEC CAUTION RESET button - Depress.
This action may reset the main and/or
standby generator.

If both MAIN GEN and STBY GEN lights remain on:

PX III With a main, standby, and EPU
generator failure, OBOGS and the OXY
LOW warning light are inoperative.
Activate EOS if above 10,000 feetcockpit altitude.

1.

Go to MAIN AND STANDBY GENERATORFAILURE (IN FLIGHT), this section.

FWhen the FLCS is powered by theaircraft battery, remain alert for
degraded flight controls. At the firstindication of degraded response, re
duce airspeed and climb to safe
ejection altitude. Eject prior to complete loss of control.

8.

HOOK switch - DN.

9.

PX II CNI, PX III C & I knob - BACKUP.

10. Minimize UHF transmissions.

T.O. GR1F16CJ1
If conditions permit:

EPU Malfunctions

11.

UNCOMMANDED EPU OPERATIONLand as soon as possible.
Fly airspeed for 11 degrees AOA approach
using fuel state when power was lost.

12. LG handle - DN. (Use DN LOCK REL button.)

If LG handle does not lower, selectBRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND. PW 229
Nozzle remains closed resulting inhigher than normal landing thrust.
13. ALT GEAR handle - Pull (190 knots maximum).

S Alternate LG extension can be used up to 300

Failures can occur which allow engine bleed air tospin the EPU turbine even though the EPU has notbeen commanded on. This may not be apparent if thethrust level and amount of bleed air are such that the
EPU is turning above or below the speed range whichturns on the EPU run light. During uncommanded
EPU operation on bleed air, EPU speed varies directlywith throttle position. High thrust settings are likely
to result in EPU failure.
The EPU may also activate for reasons not apparent to
the pilot. Although this is not an EPU malfunction, it
may be interpreted as uncommanded EPU operation.
If uncommanded EPU operation on bleed air issuspected:
NOTE

knots; however, the NLG may not fully extend
until 190 knots. Time above 190 knots should
be minimized in case there is a leak in thepneumatic lines.

The nonessential dc buses and essential dc bus may lose power. If so, thisresults in loss of power to fuel boost
and transfer pumps, PX II VHF radio,
CARA, ECM, and FCR and power for
normal weapon arming/release in
cluding selective jettison.

S WHEELS down lights and TO/LDG CONFIG
warning light function are inoperative.
Monitor LG handle warning light to verify
that LG is down.

FNWS is not available following alter

nate LG extension.

FDo not depress the ALT GEAR reset

button while pulling the ALT GEAR
handle. This action may preclude
successful LG extension.

1.

Throttle - Minimum practical thrust.

2.

Stores - Jettison (if required).
Only if required to maintain low thrust.

3.

Land as soon as possible.

If EPU is running with normal indications:
NOTE
FThe nonessential dc buses and essential

FPulling the ALT GEAR handle withdc bus lose power. This results in loss of
power to fuel boost and transfer pumps,
PX II VHF radio, CARA, ECM, and FCR
and power for normal weapon arming/
release including selective jettison.

normal system B hydraulic pressure
may result in system B hydraulicfailure within 15 minutes.

14. Consider an approachend arrestment, if
conditions permit. (Refer to CABLE ARREST
MENT, this section.)FIf the affected systems are required

for the safe recovery of the aircraft,
consider delaying/terminating EPU
operation until the systems are nolonger required.

15. Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.
After landing:
16. Stop straight ahead and have chocks installed(or engage parking brake).
17. MAIN PWR switch - MAIN PWR (until chocks
are installed).

4.

EPU - Leave running.

5.

Land as soon as possible.

6.

Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.
347

T.O. GR1F16CJ1ABNORMAL EPU OPERATION
Abnormal EPU operation after a normal startcommand is indicated by one or more of thefollowing:EPU run light flashes, indicating EPU
operation in tertiary speed control mode; EPU run
light does not come on or goes off after initial
illumination, indicating sustained underspeed oroverspeed operation; EPU HYDRAZN light does not
go off or EPU fuel quantity continues to deplete when
the throttle is advanced to assure adequate bleed air,indicating an EPU bleed air or fuel control systemproblem.
When tertiary speed control is functioning, the EPU
run light alternately cycles on and off (one to threetimes per second) as a function of EPU speed
fluctuating between the normal and slightly above
normal speed ranges. The MASTER CAUTION and
the EPU HYDRAZN lights may illuminate. Hydra
zine use occurs regardless of available engine bleed
air. The hydrazine supply depletes in approximately
10 minutes. After hydrazine depletion, the EPU
continues to operate with available bleed air on
tertiary speed control.

If EPU was turned on for an ACFT BATT FAIL or an
FLCS RLY light:
1.

EPU switch - OFF, then NORM.

2.

Land as soon as possible.

3.

Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.

If EPU was activated for other reasons:1.

Throttle - As required ( PW 229 7580, GE 129
8290 percent rpm).
Keep thrust high enough to assure adequate
bleed air if EPU fuel usage continues above
PW 229 80, GE 129 90 percent rpm or if EPU
run light is flashing. If EPU fuel is depleted
or if EPU run light goes off at low thrust, set
throttle to keep EPU run light on.

2.

EPU FUEL quantity - Monitor.

3.

Land as soon as possible.
Make an approachend arrestment, if practical, if EPU fuel depletes before landing or if
EPU run light goes off at low thrust settings.
Refer to CABLE ARRESTMENT, this section.

If tertiary speed control cannot control EPU speed(constant overspeed), total EPU failure is immi
nent. Excessive voltage generated by the EPU
generator or its PMG is regulated by theconverter/regulators before going to the FLCC.
Other electrical and avionic equipment may bedamaged by the overvoltage condition. When EPU
failure occurs, FLCS power is provided by the FLCS
PMG, main generator, standby generator, or theaircraft battery (whichever has the higher voltage).
Under some failure conditions, hydrazine may not
be available to the EPU or it may continue to deplete
even with adequate bleed air. If a hydrazine
malfunction or depletion occurs, landing must be
accomplished using an engine thrust setting
sufficient to maintain an adequate bleed air supply
to the EPU. Failure to maintain minimum engine
rpm can result in hydraulic pressure fluctuations orelectrical bus cycling. Advance throttle as required
to maintain adequate bleed air supply. (If the EPU
is the sole source of hydraulic pressure, the throttle
should be set so as to keep the EPU run light on
during mild flight control inputs.) This action may
result in a thrust level that is higher than required
for a normal straightin approach. Fully open
speedbrakes or a shallower than normal approach
may be required. A straightin approach followed by
an approachend arrestment is recommended.

348

If PTO shaft or both hydraulicsystems are failed, underspeed of the
EPU results in loss of control. Do notretard throttle completely to IDLE
until after touchdown.

If EPU underspeeds, electrical bus
cycling may affect brake operation. Fora missed engagement, attempt CHAN
1 then CHAN 2 brakes. If no braking isavailable, consider going around foranother engagement or making adepartureend arrestment. The parking brake still operates.
4.

Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.

T.O. GR1F16CJ1FLCS RLY Light
Illumination of the FLCS RLY light indicates thatvoltage on one or more of the four FLCC branches
connected to the aircraft battery is inadequate(below 20V). Therefore, should main generator,
standby generator, and FLCS PMG power be lost(engine shutdown, PTO shaft failure, ADG failure),
power would not be available to one or more FLCS
branches until the EPU is on line. The EPU must be
turned on if the FLCS RLY light cannot be
extinguished.
1.

FLCS PWR TEST switch - TEST, momentarily.

If FLCS RLY light goes off:
2.

Land as soon as practical.

If FLCS RLY light remains on:
2.

EPU switch - ON.
NOTE
FThe nonessential dc buses and essential dc bus lose power. This results inloss of power to fuel boost and transfer
pumps, PX II VHF radio, CARA, ECM,
and FCR and power for normal weapon
arming/release including selective
jettison.

FIf the affected systems are required forthe safe recovery of the aircraft,
consider delaying/terminating EPU
operation until the systems are nolonger required.

3.

Land as soon as practical.

4.

Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.

If EPU runs abnormally:
3.

EPU switch - OFF, then NORM.

4.

Land as soon as possible.

Partial Electrical Power Loss
Loss of power to several systems or indicators without
any indications on the ELEC control panel may be theresult of wire harness chafing or the loss of power toone or more ac or dc buses. Refer to EMERGENCY
POWER DISTRIBUTION, this section, to determine
affected buses and equipment. If one item on a bus ispowered, then that bus should be considered
powered. Wire harness chafing can affect numerous
items on more than one bus without causing loss of
power to a bus.
A failed overcurrent sensing contactor (OCSC) may
be the cause of the partial power loss. If the OCSC
failure involves electrical shorting, power from themain generator is degraded while the short ispresent. An obvious effect of this degraded power isloss of the HUD and MFD's. If the short quickly burns
itself open, only the system(s) which receives power
through the failed OCSC is lost. Power to thenonessential ac No. 1, emergency ac No. 1, emergency
dc No. 1, and nonessential dc buses goes through one
OCSC. Failure of this OCSC results in a loss of power
to many items, including the FCR, CADC, stick trim,
TACAN, VHF radio, DED, FUEL FLOW indicator,
primary instrument and console lights, VVI, AOA
indicator, and PX II FCC. PX III The MMC enters adegraded mode that includes loss of the HUD. Failure
effects for the other OCSC's are limited to individualstore stations, the FCR, or the nacelle ac buses. If theshort persists for 23 seconds, the main generator
goes off line and power is removed from all eightovercurrent sensing contactors. The standby genera
tor should automatically come on line to power the
essential, emergency, and battery buses. Refer to
SINGLE GENERATOR FAILURES (IN FLIGHT),
this section if the MAIN GEN FAIL light is on.
OCSC's can also fail open without electricalshorting. These failures do not result in degraded
power from the main generator but do cause loss of
power to the buses/items protected by the OCSC.
The ELEC CAUTION RESET button can be used toreset a failed open OCSC; however, the OCSC may
not remain reset if the fault persists.
The buses with a resettable overcurrent sensing
contactor are nonessential ac bus No. 1 and thenacelle nonessential ac bus.
The items with a nonresettable overcurrent sensing
contactor are the radar ac bus; stations 3, 5, and 7;
and left and right inlet stations.
1.

ELEC CAUTION RESET button-Depress.
The failed open OCSC may reset.

If power is restored:
2.

Land as soon as practical.

Change 1349

T.O. GR1F16CJ1
If power is not restored:
2.

Determine the power status of electrical buses.
Refer to EMERGENCY POWER DISTRIBUTION, this section, to determine thepower status of individual buses. If one item
on a bus is powered, then that bus should beconsidered powered.
NOTE
Determining the status of the battery
buses is critical for a safe recovery of
the aircraft.

If one or both emergency ac buses are not powered:3.

EPU switch-ON.

Aircraft Battery Failure
Aircraft battery system failure is indicated by the
ELEC SYS and ACFT BATT FAIL lights. The FLCSRLY light also illuminates if the failure involves low
battery voltage. If the battery fails and both mainand standby generators subsequently drop off line,
power is not available to activate the EPU. If thebattery fails and subsequent loss of main generator,
standby generator, and FLCS PMG power occurs(e.g., engine shutdown, PTO shaft failure, ADG
failure), power is not available to activate the EPU
or control the aircraft. Thus, it is necessary to turnthe EPU on after battery failure is indicated. The
ACFT BATT FAIL light illuminates only for low
battery voltage while in flight (approximately 20
volts). If a battery charger failure occurs in flight andis still present after landing, the ACFT BATT FAIL
light illuminates 60 seconds after WOW.

NOTE
FThe nonessential dc buses and essential dc bus lose power. This results inloss of power to fuel boost and transfer
pumps, PX II VHF radio, CARA, ECM,
and FCR and power for normal weapon
arming/release including selective jet
tison.

FIf the affected systems are required forthe safe recovery of the aircraft,
consider delaying/terminating EPU
operation until the systems are nolonger required.

If the battery buses PX III and emergency dc bus
No. 2 are not powered,:4.

Consider a net arrestment, refer to NET
ARRESTMENT, this section.

If net arrestment is not available:
5.

Consider a gear up landing, refer to LANDING
WITH LG UNSAFE/UP, this section.
If power to the battery buses is lost after thelanding gear has been extended, the landing
gear cannot be raised.

6.

Refer to EMERGENCY POWER DISTRIBU
TION, this section.

7.

Land as soon as possible.

If EPU was activated:
8.

Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.

350Change 1

If the aircraft battery fails (and EPU isoff), do not taxi except to clear runway.
Subsequent loss of the main andstandby generators results in loss of all
braking, NWS, hook, radios, and dragchute.
1.

EPU switch-ON.
NOTE
FThe nonessential dc buses and essential dc bus lose power. This results inloss of power to fuel boost and transfer
pumps, PX II VHF radio, CARA, ECM,
and FCR and power for normal weapon
arming/release including selective jet
tison.

FIf the affected systems are required forthe safe recovery of the aircraft,
consider delaying/terminating EPU
operation until the systems are nolonger required.

FIf both radios become inoperative afteran aircraft battery failure indication,refer to PARTIAL ELECTRICAL
POWER LOSS, this section.

FThe ACFT BATT FAIL light may

subsequently extinguish. This should
not be interpreted to mean that thebattery has recharged. It may indicatethat the battery voltage is so low thatthe light cannot remain illuminated.

T.O. GR1F16CJ1
2.

Land as soon as practical.

6.

3.

Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.

Prior to shutdown:

If EPU runs abnormally:
4.

EPU switch-OFF, then NORM.

5.

Land as soon as possible.

Refer to ACTIVATED
LEAK, this section.

7.

Loose items-Secure.

8.

Canopy-Open.

EPU/HYDRAZINE

Emergency Power Distribution
Refer to figure 36.

351

T.O. GR1F16CJ1Emergency Power Distribution
MAIN GENERATOR FAILED
BUS ASSIGNMENTSYSTEM

INOPERATIVE EQUIPMENTNONESS AC
NO. 1

NO. 2

ENGINE

GE129
EMSC

NAV/COMM

PX II

FUEL

Pumps 1, 2, 4 & 5

X

CFT PumpsX

PX IIINACELLENONESS DC
NO. 1

NO. 2
X

VHF Radio

*
*
X

GE129
FUEL/OIL
HOT Caution Light(oil hot signal)

STORES MGTAVIONICS

AIM120
Stations 3, 5 & 7-EO, RadarGuided
Weapons, ECM Pods*****

Stations 4 & 6-EO, RadarGuided Weapons

X

DTU

X

PX II

FCC

X

FCR
PX II

Radar
GPS

X

TWS
LIGHTS

OTHER

*

X

X

Flood Console

X

Flood Instrument

X

Formation

X

Taxi

X

PX IIIDASIU*ECM Control*

Halon Heater
PX III

X

HMCS

X

Inlet Strut Heater
PX II

X

INS Heater

X

Nacelle Ejector Shutoff

X

Seat Adjustment

X

Total Temperature Probe Heater

X

NOTE:
Equipment on nonessential ac bus No. 1 or nonessential ac bus No. 2 may be functional with MAIN GEN
light on (bus contactor failure).
*Aft equipment bay nonessential dc bus.
**Overcurrent sensing contactors.
***Nacelle nonessential ac bus.
Figure 36.(Sheet 1)352

T.O. GR1F16CJ1Emergency Power Distribution
MAIN AND STANDBY GENERATORS FAILED(All equipment from sheet 1 plus the following:)

SYSTEMFUELINOPERATIVE EQUIPMENT
Pump 3

BUS ASSIGNMENTESS ACESS DCX

X

Tank Inerting
STORES MGTAVIONICS

OTHER

X

AIM9

*

Arm and Release Power-Stations 1 Thru 9X

Radar Altimeter

X

MFD's

X

PFLD

*

Air Data Probe Heater (fuselage)*

PX III

D

ASHM

XBattery ChargerX

PX III Data Link

X

NOTE:
Equipment on this sheet may operate if MAIN GEN light was caused by bus contactor failure at nonessential bus No. 1.
*Nacelle essential ac bus.

Figure 36.(Sheet 2)

353

T.O. GR1F16CJ1Emergency Power Distribution
MAIN, STANDBY, AND EPU GENERATORS FAILED(All equipment from sheets 1 and 2 plus the following:)
BUS ASSIGNMENTSYSTEM

INOPERATIVE EQUIPMENT

EMER AC
NO. 1

ENGINE

FLIGHT
INSTRUMENT

FUEL

FLIGHT
CONTROLSEngine ANTI ICE Switch
ENGINE FAULT Caution Light
Engine Ice Detector
Fire/Overheat Detect and Test
HYD PRESS Indicators
GE129
Low
Energy Ignition PowerNOZ POS Indicator
OIL Pressure Indicator
ADI
Altimeter (ELECT)AOA Indexer
AOA Indicator
HSI
Turn Needle
INSTR MODE Select Switch
VVI
Automatic Forward Fuel Transfer
FUEL FLOW Indicator
FUEL LOW Caution Lights
FUEL Quantity Indicator
Autopilot
DBU ON Warning Light (branches A & B)
DBU ON Warning Light (branches C & D)FLCS
FAULT Caution LightC DF(branches A & B)

NO. 2

EMER DC
NO. 1

X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X

DR
FLCS FAULT Caution Light(branches C & D)
FLCS RESET Switch (branches A & B)
FLCS RESET Switch (branches C & D)
FLCS Power Source (branches A & B)
FLCS Power Source (branches C & D)
FLCS Warning Light (branches A & B)
FLCS Warning Light (branches C & D)

Figure 36.(Sheet 3)354

NO. 2

X
X
X
X
X
X
X

T.O. GR1F16CJ1Emergency Power Distribution
MAIN, STANDBY, AND EPU GENERATORS FAILED-CONT(All equipment from sheets 1, 2, and 3 plus the following:)
BUS ASSIGNMENTSYSTEM

INOPERATIVE EQUIPMENT

EMER AC
NO. 1

FLIGHT
CONTROLS(cont)LEF's

NAV/COMM

PX III

NO. 2

EMER DC
NO. 1

X

Speedbrakes

X

Stick Trim

X

EGI

X

IFF
PX II

X
X

INS

X

ILS

X

TACAN
PX IIISTORES MGTX

XX

VHF RadioCALT
REL Button

X

CIU*

X

Chaff Dispenser

X
X

Gun

X

X

EMER JETT Button*

X

MASTER ARM SwitchX

MSL STEP Switch

X

X

NUCLEAR CONSENT SwitchX

STORES CONFIG Caution Light

X

C
WPN
REL Button
DF XWPN REL Button
DR
AVIONICS

NO. 2

X

CADC

X

CADC Caution Light

X

HUD

X

HUD/CTVSX

ICP/IKPX

MFD Video ControlX

PX III

MMC*

X

Upfront Controls

X
X

X

X
X

* Indicates redundancy.

Figure 36.(Sheet 4)

355

T.O. GR1F16CJ1Emergency Power Distribution
MAIN, STANDBY, AND EPU GENERATORS FAILED-CONT(All equipment from sheets 1, 2, 3, and 4 plus the following)
BUS ASSIGNMENTSYSTEM

INOPERATIVE EQUIPMENT

EMER AC
NO. 1

LIGHTSNO. 2

ANTICOLLISION StrobeX

AR (flood)

X

EMER DCNO. 1

AR (slipway)

X

LandingX

LANDING/TAXI/External SwitchesX

MAL & IND LTS TEST/BRT DIMXPOSITION

LG/NWS/BRAKESX

PRIMARY CONSOLESXPRIMARY INST PNLXLG Hydraulic IsolationX

LG Sequence (doors)X

LG UPDN Command

OTHER

X

NWS

X

WHEELS Down Lights

X

Air Data Probe Heater (nose)X

AOA Probe HeatersX

AR SystemX

AVTR/CTVSXCABIN PRESS Caution LightXCAMERA/GUN Trigger

X

Cockpit Pressure Dump Capability

X

Cockpit Temperature Control

X

Engine Bleed Air Valves (close capability)

X

EQUIP HOT Caution LightX

INLET ICING Caution LightX

PX II LIQUIDOXYGEN Quantity
C DF
Indicator

Figure 36.(Sheet 5)356

NO. 2

X

T.O. GR1F16CJ1Emergency Power Distribution
MAIN, STANDBY, AND EPU GENERATORS FAILED-CONT(All equipment from sheets 1, 2, 3, 4, and 5 plus the following)
BUS ASSIGNMENTSYSTEM

INOPERATIVE EQUIPMENT

EMER ACNO. 1

OTHER continued

NO. 2

EMER DC
NO. 1

NO. 2

PX II

OXY LOW Caution LightX

PX IIIOXY LOW Warning LightX

PX IIIOBOGS Caution LightX

PX IIIOBOGS Concentrator

PX IIIOBOGS Monitor

X
X

Probe Heat Monitor

X

PROBE HEAT Switch

X

SEAT NOT ARMED Caution Light

X

Figure 36.(Sheet 6)

357

T.O. GR1F16CJ1Emergency Power Distribution
OPERATING EQUIPMENT-MAIN, STANDBY, AND EPU GENERATORS FAILED
BUS ASSIGNMENTSYSTEM
ENGINEOPERATING EQUIPMENT

BATTERY
NO. 1

BATTERY
NO. 2

PW229

EDU

X

GE129

Electrical Throttle PositionX

PRI (no supersonic stall protection)*
PRI/SEC Transfer Circuit*
INSTRUMENTS

Airspeed/Mach Indicator*
Altimeter (PNEU)*
FTIT IndicatorX

RPM Indicator

X

SAI
FUEL

X

External Fuel Transfer*
FUEL MASTER Switch

X

FFP*
FLIGHT
CONTROLS

Functional (except LEF's, speedbrakes, autopilot, andstick trim)*

NAV/COMMIntercom

X

Magnetic Compass*
LIGHTS
LG/NWS/BRAKES

UHF RadioX

Spotlights

X

Utility Light

X

Alternate LG Extension*
Antiskid/Channel 1 Brakes

X

Antiskid/Channel 2 Brakes

X

LG Uplock/Downlock

X

MLG WOW (branches A & B)X

MLG WOW (branches C & D)
NLG WOW (branches A & B)X

NLG WOW (branches C & D)

X

Parking Brake

X*Indicates items that do not require power through the battery buses.
Figure 36.(Sheet 7)358

X

T.O. GR1F16CJ1Emergency Power Distribution
OPERATING EQUIPMENT - MAIN, STANDBY, AND EPU GENERATORSFAILED - CONTBUS ASSIGNMENTSYSTEMWARNING
LIGHTSCAUTION
LIGHTSOPERATING EQUIPMENT

BATTERY
NO. 1

CANOPYX

ENGINEX

HYD/OIL PRESSX

LG Warning (handle)

X

ANTI SKIDXELEC SYS

X

HOOK

X

MASTER CAUTION

X

SEC
OTHER

BATTERY
NO. 2

X

Canopy Activation*
Drag Chute

X

EPU

X

Hook

X
X

JFS

X

MAIN PWR Switch

X

VMS

X

*Indicates items that do not require power through the battery buses.

Figure 36.(Sheet 8)

359

T.O. GR1F16CJ1Emergency Power Distribution
PARTIAL ELECTRICAL POWER LOSSNONESSENTIAL AC BUS NO. 1
PX II FCC
Seat AdjustmentTWS
PX III HMCS

MAIN
GEN

NONESSENTIAL AC BUS NO. 2
FLOOD Instrument/
Console Lights

STBY
GEN

ESSENTIAL AC BUS
MFD's (off)

EMERGENCY AC BUS NO. 1

EPU
GEN

EMERGENCY AC BUS NO. 2

AOA Indicator (off)FUEL FLOW (frozen)
VVI (off)
Critical items lost:
D CADCD PX II INS (6090 sec
of use)HSI (off flag)
HYD PRESS (frozen)
OIL Pressure (frozen)
Critical items lost:
D ADI, HSI, Fuel Quantity
D TACANEMERGENCY DC BUS NO. 1

EMERGENCY DC BUS NO. 2

AOA Indexer
Speedbrakes
PX III VHF Radio
Critical items lost:
D LG DOWN Permission
Button (DN LOCK REL
button required)
D LG WHEELS Down Lights
D TACAND PX III EGI

Autopilot
HUD
SEAT NOT ARMED Caution Light
Critical items lost:
D LGUPDN Command(ALT GEAR extension
required)
D ILS

BATTERY BUS NO. 1

AIRCRAFT BATTERYBATTERY BUS NO. 2
EPU Fuel (indicates < zero)
SAI (off flag)
Critical items lost:
D PX II Channel 2 Brakes
and Parking Brake
D PX III Parking Brake
D Hook
D LGWarning Light(handle)FTIT (indicates < 200)
Intercom*
RPM (indicates < zero)
UHF Radio
Critical items lost:
D Channel 1 Brakes
D Drag Chute
D LG WHEELS Down Lights*VHF radio is also inoperative because
the ability to key either radio is lost.

GR1F16CJ12120X37

Figure 36.(Sheet 9)360Change 1

T.O. GR1F16CJ1
ENGINE MALFUNCTIONS PW 229
The EDU compares expected versus actual engine
operation. The purpose of the EDU MFL is to provide
maintenance personnel with an early indication of anengine condition which requires correction. No action isrequired for an engine MFL at anytime during a flight.
Low altitude, for engine malfunction purposes, isgenerally defined as 10,000 feet AGL or below. If an
engine malfunction is suspected, the initial reaction
should be to trade excess airspeed for altitude. Unlessa suitable airfield is within gliding distance, turns
should be avoided as they decrease the amount oftime/altitude available to successfully recover engine
performance or prepare for ejection. Optimizing theexchange of airspeed for altitude must be a priority
action for any engine malfunction. Above 350 knots,
more time is available by performing a zoom climb
using a 3g pullup to 30degree climb until
approaching the desired airspeed (use approximately
50 knots lead point) and then initiating a zerogpushover. Below 350 knots and above the minimum
recommended ejection altitude, more time isavailable by performing a constant altitude deceleration to the desired airspeed. If below the minimum
recommended ejection altitude and below 350 knots,
primary concern should be to trade excess airspeed
for altitude in preparation for ejection. If appropriate,
jettison stores as soon as possible.

With engine failure or flameout,
OBOGS is inoperative. Activate EOS if
OXY LOW warning light illuminates
above 10,000 feet cockpit altitude.
For any situation where automatic activation of the
EPU is relied upon, verify that the EPU run light is
on to insure that the EPU has started. If the EPU run
light is off, position the EPU switch to ON.
Idle thrust in SEC during ground operation isapproximately twice that in PRI. After landing in SEC,
consider minimizing taxi distance and consider
following HOT BRAKES procedures, this section.
Engine Fire PW 229
Generally, the first indication of fire in the engine
compartment is the ENG FIRE warning light.
Abnormal fuel indications (quantity/flow) may also bepresent. FTIT probably will not be higher than normal.
Explosions, vibrations, or engine instrument fluctua
tions are usually indicative of a serious engine problem;
engine failure may be imminent. Immediate action

should be taken to reduce thrust to the minimum
practical level after attaining safe ejection parameters.
If within gliding distance of a suitable runway, consider
shutting the engine down. Sufficient time should exist
to analyze the situation and make an ejection versus
land decision. The ejection decision should be based on
visual and/or cockpit indications that the fire is
persisting. Cockpit indications include continued
illumination of the ENG FIRE warning light andsubsequent FLCS malfunctions/degraded flight con
trols or subsequent loss of either hydraulic system.
Fires can also occur in the exhaust nozzle area whenusing AB. These fires are the result of portions of thenozzle failing which allows the AB plume to burnthrough the nozzle. Ventilation should inhibit forward
movement of the fire into and through the engine bay.
Since these fires are aft of the detection circuit, the
ENG FIRE warning light will not illuminate.
Additionally, the nozzle position indications arenormal, and there are no vibrations or instrument
fluctuations. In most cases, these ABrelated nozzle
fires are detected by someone outside the aircraft(wingman, tower, etc.). When operating in AB and afire is reported at the rear of the aircraft, retard
throttle below AB immediately. This action should
extinguish a nozzle fire within approximately 3045
seconds and minimize damage to the aircraft skin,speedbrakes, nozzle, and flight controls; however,nozzle damage may result in a noticeable thrust loss.
A failure that causes an oil leak may also result in an
oilfed fire in the AB section. The fire may continue forseveral minutes after the engine fails or is shut down
(until the oil supply is exhausted). Since the fire is
likely to be contained within the engine, the ENG
FIRE warning light does not illuminate.
If on takeoff and conditions permit:
1.

Abort.

If takeoff is continued:
1.

Climb.
Maintain takeoff thrust until minimum
recommended ejection altitude is attained
and then throttle to minimum practical.

2.

Stores-Jettison (if required).

At a safe altitude:3.

Throttle-Minimum practical.
If fire occurred in AB, ENG FIRE warning
light may not illuminate. Fire should
extinguish after throttle is retarded; how
ever, nozzle damage may result in lower thannormal thrust.
361

T.O. GR1F16CJ1
If ENG FIRE warning light goes off:
4.

FIRE & OHEAT DETECT button-Depress.
Determine if fire detection circuit is func
tional.

F With the ECS shut down or the AIR

SOURCE knob in OFF or RAM, the gsuit
does not inflate and PBG is disabled.

If fire persists:
5.

FWith the AIR SOURCE knob in OFF or

RAM, OBOGS is inoperative. Activate
EOS if OXY LOW warning lightilluminates above 10,000 feet cockpit
altitude.

Eject.

If fire indications cease:
5.

Land as soon as possible.

OVERHEAT Caution Light PW 229
Detection of an overheat condition in the engine
compartment, ECS bay, MLG wheel wells, or EPU
bay illuminates the OVERHEAT caution light.
Accomplish as many of the following as required toextinguish the light. If the light goes off, verify the
integrity of the detection circuit by depressing the
FIRE & OHEAT DETECT button and land as soon as
possible.

6.

Descend to below 25,000 feet and reduce airspeed to below 500 knots.

When airspeed
depressurized:7.

isreducedandcockpitis

AIR SOURCE knob-RAM (below 25,000 feet).
External fuel cannot be transferred in OFF or
RAM. Consider jettisoning tanks to decreasedrag if range is critical and the ECS cannot be
turned on for short periods of time to transfer
fuel.

If OVERHEAT caution light illuminates:
1.

Throttle-Minimum practical.

2.

FIRE & OHEAT DETECT button-Depress.
Determine if fire detection circuit is func
tional.

FWith the ECS shut down or the AIR

SOURCE knob in OFF or RAM, thegsuit does not inflate and PBG isdisabled.

FWith the AIR SOURCE knob in OFF or

If OVERHEAT caution light remains on (or detectcircuit checks bad) and EPU is running:
3.

EPU switch-OFF (if feasible).
If the EPU was manually turned on, consider
turning it off to determine if it is the source ofthe overheat condition. If the OVERHEAT
caution light remains on, the EPU should beturned back on.

RAM, OBOGS is inoperative. Activate
EOS if OXY LOW warning lightilluminates above 10,000 feet cockpit
altitude.

8.

NOTE
If in VMC and the ADI and HSI are notrequired for flight, the EGI should beconsidered nonessential.

If OVERHEAT caution light remains on (or detectcircuit checks bad):
4.

OXYGEN-100%.

5.

AIR SOURCE knob-OFF.
External fuel cannot be transferred in OFF or
RAM. Consider jettisoning tanks to decreasedrag if range is critical and the ECS cannot be
turned on for short periods of time to transfer
fuel.

362

Nonessential electrical equipment-Off.

If OVERHEAT caution light still remains on (ordetect circuit checks bad):
9.

TANK INERTING switch-TANK INERTING even if Halon is not available.

10. LG handle-DN (300 knots/0.65 mach maximum). (Use DN LOCK REL button if required.)T.O. GR1F16CJ1

If LG handle does not lower, selectBRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND. Nozzle
remains closed, resulting in higher
than normal landing thrust.
11.

Land as soon as possible.

Engine Vibrations PW 229
Some engines exhibit low frequency vibrations which
are nondamaging to both the airframe and engine.
The vibrations should disappear if engine rpm iseither increased or decreased.
Vibrations that change in intensity with throttle
movement and are present across the throttle/rpm
range may indicate a potential engine malfunction.
If vibrations persist:1.

Throttle-Minimum practical.

2.

Land as soon as possible.

zero oil pressure, any throttle movement may cause
the engine to seize. Minimize maneuvering g tominimize loads. Plan an approach which allows aflameout landing from any position should engine
seize. Refer to SIMULATED FLAMEOUT (SFO)
LANDING and FLAMEOUT LANDING, this section.
The EPU should be manually activated; otherwise, if
the EPU does not start automatically when theengine seizes, the short time remaining before loss ofcontrol may be inadequate for recognition of the EPU
failure and corrective action. Monitor hydrazine use
after activating the EPU. If consumption rate is too
high, cycle EPU switch to OFF, then NORM toconserve hydrazine. Be prepared to place the EPU
switch back to ON if the engine seizes.
If an oil pressure malfunction is suspected:
1.

Attain desired cruise altitude.
The rate of oil loss is decreased at low
altitudes and low throttle settings.

2.

Stores-Jettison (if required).

3.

Throttle-Approximately 80 percent rpm.

4.

EPU switch-ON.
Monitor hydrazine use. If consumption rate is
too high, cycle EPU switch to OFF, thenNORM to conserve hydrazine. Be prepared toplace EPU switch back to ON if the engine
seizes.

5.

Throttle-Do not move until landing is assured.

Oil System Malfunction PW 229
An oil system malfunction is characterized by apressure (including fluctuations) below 15 psi at
IDLE or 30 psi at MIL, a pressure above 95 psi at any
thrust setting, pressure fluctuations greater than "5
psi at IDLE or "10 psi above IDLE, or by a lack of oilpressure rise when the throttle is advanced. The OIL
pressure indicator can be used as an early indicationof oil loss. An indication of excessive oil loss is the lackof oil pressure rise when the throttle is advanced in
the IDLE to MIL range. These conditions may notoccur until approximately onehalf the usable oil islost. The HYD/OIL PRESS warning light may notilluminate until most of the usable oil is lost. At thefirst indication of an oil system malfunction, take
immediate action to land as soon as possible.
Climbing to a higher altitude allows higher cruise
airspeed and increases glide range. However, if the oil
malfunction is caused by an internal engine oil leak,the rate of oil loss is decreased at low altitude andthrottle settings. Usually it is advisable to climb to areasonable cruise altitude. Once at altitude, retard
throttle to approximately 80 percent rpm and do not
move the throttle unless absolutely required. WithFThrottle movement/rpm change may

cause engine seizure.

FDo not start the JFS if engine seizure

has occurred or is anticipated. Start
ing the JFS may result in no brake/JFS
accumulator pressure for the brakes.

6.

Land as soon as possible.
Plan to fly an SFO. Refer to SIMULATEDFLAMEOUT (SFO) LANDING and FLAMEOUT LANDING, this section.

7.

Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.

363

T.O. GR1F16CJ1NOTE

ENGINE FAULT Caution Light PW 229

FCompressor stalls may result from

Illumination of the ENGINE FAULT caution lightindicates that an engine PFL item was detected.

throttle movement in SEC above
40,000 feet MSL.

If ENGINE FAULT caution light illuminates:
1.

40,000 feet MSL may result in stalls.

PFLD-Note PFL(s) displayed.
NOTE
If ENG BUS FAIL PFL is displayed or
has been displayed, MUX communica
tion with the EDU is no longer possible.
Subsequently, if an engine PFL occurs,
the ENGINE FAULT caution lightilluminates but cannot be reset and thatPFL cannot be displayed on the PFLD.

2.

FTransfers from SEC to PRI above

C DF FACK, DR FAULT ACK button-Depressto acknowledge fault.

Automatic transfers to SEC after an engine
alternator failure may also cause engine stalls. The
combination of stalls and an erroneously low rpm
indication may be incorrectly interpreted as anonrecoverable stall. Confirm that a nonrecoverable
stall actually exists before shutting down the engine.
If the SEC caution light illuminates while supersonic,
do not retard throttle below MIL until subsonic.
During landing in SEC, idle thrust is approximately
twice that in PRI with a normal nozzle because thenozzle is closed.
NOTE

If ENGINE FAULT caution light does not reset when
the fault is acknowledged:3.

Throttle-85 percent rpm or less.

4.

Land as soon as possible.

If ENGINE FAULT caution light resets when thefault is acknowledged:
3.

Refer to PILOT FAULT LIST-ENGINE, thissection.

4.

AB RESET switch-AB RESET, thenNORM.
This action resets the DEEC and may clear
the failure condition.

5.

FACK, DR FAULT ACK button-Depress to perform fault recall.
The failure condition no longer exists if the
PFL is not present during the fault recall.

CC

The ENG CONT switch should not be
returned to C DF PRI, DR NORM after
landing in an attempt to open thenozzle and decrease thrust.
If SEC caution light illuminates while supersonic:
1.

Retarding the throttle below MIL
while supersonic may induce inlet
buzz which produces severe cockpit
vibration and probable engine stalls.

DF

DF

SEC Caution Light PW 229
Illumination of the SEC caution light indicates thatthe engine is operating in SEC. If the ENG CONT
switch is in C DF PRI, DR NORM and the SEC
caution light is on, an automatic transfer to SEC has
occurred. The transfer may be due to a DEEC
malfunction, the DEEC sensing the loss of a critical
input signal to the DEEC, or loss of power to the
DEEC (engine alternator failure).
364Change 1

Throttle-Do not retard below MIL until
subsonic.

When subsonic or if SEC caution light illuminates
while subsonic:
2.

Throttle-Verify engine responds normally tothrottle movement from IDLE to MIL; set as
required.
AB operation is inhibited. Above 40,000 feet
MSL, minimize throttle movement.

If the rpm indication is also zero or
erroneously low, the engine alternatormay have failed. If the engine is shut
down, an airstart may not be possible.

T.O. GR1F16CJ1
3.

ENG CONT switch-SEC.

Abnormal Engine Response PW 229

4.

Land as soon as practical.
During landing in SEC, idle thrust isapproximately twice that in PRI with anormal nozzle.

Refer to LOW THRUST ON TAKEOFF OR AT LOW
ALTITUDE (NONAB) PW 229 , this section, if
appropriate.

If engine is operating abnormally in SEC:
5.

Refer to ABNORMAL ENGINE RESPONSE
PW 229 , this section.

FTIT Indicator Failure PW 229
Certain failures of the FTIT indicator can cause
erroneous indications above 1100_C and illuminationof the ENGINE warning light. If all other engine
indications are normal and the engine responds
normally to throttle movement, the engine should not
be shut down. Routine missions should not be
continued since FTIT cannot be monitored.
Zero RPM/Erroneous RPM Indication PW 229
If the RPM indicator displays a zero or erroneous
indication while other engine instruments indicate
normal operation, the cause is loss of power to theindicator or indicator failure. Loss of power to theindicator causes the ENGINE warning light toilluminate. RPM indicator failure may not cause
ENGINE warning light illumination.

Assume engine alternator is inopera
tive or malfunctioning. If the engine isshut down, an airstart may not be
possible.
If the RPM indicator displays a zero or erroneously
low indication accompanied by an automatic transfer
to SEC (SEC caution light illuminated), the engine
alternator may have failed. Since rpm cannot bemonitored, routine missions should not be continued.
If SEC caution light is illuminated:
1.

Go to SEC CAUTION LIGHT PW 229 , thissection.

If SEC caution light is not illuminated:
1.

Land as soon as practical.

Abnormal engine response is varied and generallyindicated by abnormal thrust in relation to throttle
position, engine oscillations (either continuous,
momentary, or recurring), a complete lack of engine
response to throttle movement, autoacceleration/de
celeration, exhaust nozzle failure, or insufficientthrust. The DEEC detects and automaticallyattempts to take corrective action for engine
malfunctions. This action may result in partially or
totally inhibited AB operation or in engine control
being transferred to SEC. The action taken by the
DEEC may be indicated by illumination of either orboth the ENGINE FAULT and SEC caution lights.
If thrust is too low to sustain level flight, turnimmediately toward the nearest suitable runway andestablish 250 knots airspeed. Consider jettisoning
stores to increase flying time available to complete
actions designed to restore usable thrust and improverange in the event those actions are unsuccessful.
NOTE
In a partial thrust situation, thrustavailable may increase as altitudedecreases. 250 knots approximates theairspeed at which thrust required forlevel flight is the lowest.
A throttle linkage problem should be suspected if
throttle movements in both PRI and SEC produce
either no rpm change or an rpm increase but no rpm
decrease. In either case, the OFF position does notshut down the engine. If the throttle is stuck andthrust is suitable for sustained flight, attempts to free
the throttle should be delayed until within gliding
distance of a suitable landing field. If throttle is stuckin AB, placing the ENG CONT switch to SEC
terminates AB and provides SEC MIL thrust. If
throttle is stuck or otherwise prevented from normal
movement, control might be regained by depressing
the cutoff release, rotating the throttle outboard,and applying necessary force.
The engine may roll back, which prevents the engine
from reaching normal rpm and FTIT levels when thethrottle is advanced. This rollback is generally causedby the DEEC sensing an outoflimits condition andmay not be accompanied by a SEC caution light. For
this situation or any abnormal engine response below
AB, follow the procedures of this section until thesituation is corrected.
365

T.O. GR1F16CJ1
If thrust is too high to permit a safe landing, use
excess thrust to climb and maneuver toward thenearest suitable airfield. Once high key for a flameout
landing is assured, follow procedures as outlined in
FLAMEOUT LANDING, this section. Activate the
JFS and EPU and then shut down the engine as soon
as landing is assured (normally high key) by placing
the throttle to OFF or, if necessary, by placing the
FUEL MASTER switch to OFF.
If the decision is made to manually select SEC while
subsonic below 40,000 feet MSL, set the throttle tomidrange or above before positioning the ENG CONT
switch to SEC.

If thrust is low and nozzle is suspected to be failed
open, damaged, or missing:
2.

FJettison stores when necessary to

C DF AB RESET switch-AB RESET, then
NORM.

4.

Airspeed-250 knots (if thrust is too low tosustain level flight).

If problem still exists:
5.

Throttle-IDLE.

6.

ANTI ICE switch-OFF.
Stalls may be caused by the antiice valve
failing to close at high throttle settings(above midrange).

7.

Throttle-Slowly advance to minimum practical.
Attempts to establish a minimum practicalthrottle setting that provides sufficientthrust may result in repeated stalls thatclear when the throttle is retarded. Notestalled RPM/throttle position and attempt
to establish a lower throttle setting thatprovides sufficient thrust.

If current thrust will allow a safe landing:increase flying time available toaccomplish actions designed to restore
thrust.

8.

NOTE

8.

Throttle-Midrange.

FTransfer to SEC removes stall recovery

9.

ENG CONT switch-SEC.
Transfer to SEC while supersonic should beaccomplished with the throttle at MIL; if thethrottle can not be retarded to MIL, transfer
to SEC is permissible with the throttle in AB.
Subsonic transfers to SEC below 40,000 feet
MSL should be accomplished with thethrottle at midrange or above.

logic. If SEC is selected while theengine is stalling, a stagnation may
occur.

returned to C DF PRI, DR NORM after
landing in an attempt to open thenozzle and decrease thrust.

If in AB or supersonic:
Throttle-MIL.

Retarding the throttle below MIL
while supersonic may induce inlet
buzz which produces severe cockpit
vibration and probable engine stalls.
366Change 1

Land as soon as possible.

If suitable thrust cannot be attained or thrust is toohigh to permit a safe landing:FThe ENG CONT switch should not be

1.

this

3.

FIf the throttle is stuck and thrust issuitable for sustained flight, attempts
to free the throttle should be delayed
until within gliding distance of asuitable landing field.

PW 229 ,If problem still exists:

FFailure to monitor sink rate and height

above terrain while applying low
thrust recovery procedures can resultin ejection outside ejection seat perfor
mance envelope.

Refer to NOZZLE FAILURE
section.

10. Throttle-Minimum practical.
If current SEC thrust will allow a safe landing:11.

Land as soon as practical.
During landing in SEC, idle thrust isapproximately twice that in PRI.

An SFO is not recommended if engine
is operating satisfactorily in SEC.

T.O. GR1F16CJ1When landing is assured:

13. Hook switch-DN (if required).

12. Throttle - Verify engine responds normally tothrottle movement from IDLE to MIL; set as
required.
If suitable thrust cannot be attained:11.

ENG CONT switch- C

DF PRI, DR NORM.

12. Throttle-AB (if required to sustain level
flight).
13. Land as soon as possible.
If thrust is too high to permit a safe landing:
NOTE
If throttle is stuck, control might beregained by depressing the cutoff
release, rotating the throttle outboard,and applying necessary force.
11.

Plan a flameout landing. Refer to FLAMEOUTLANDING, this section.

Do not start the JFS if engine seizure
has occurred or is anticipated or ifengine failure is a result of fuelstarvation. Starting the JFS may
result in no brake/JFS accumulator
pressure for the brakes.
When prepared to land (normally high key):

Delaying engine shutdown can resultin a long, fast landing. Wheel braking
is less effective due to lack of WOW andthere is an increased probability of a
missed cable engagement.
12. Throttle-OFF.
If throttle is stuck or engine does not respond,
shut down the engine with the FUELMASTER switch. At MIL, the engine flamesout in approximately 6 seconds. At IDLE, theengine flames out in approximately 45
seconds.

The hook may miss the cable if theaircraft is not slow enough to compress
the MLG struts sufficiently to makeWOW or if forward stick pressure is held.
Nozzle Failure PW 229
Exhaust nozzle malfunctions and nozzle control
system malfunctions can result in the nozzle being too
far open or too far closed. These malfunctions canresult in loss of AB capability, engine stalls, or low
thrust. Separation of the nozzle assembly from theengine is also possible and results in low thrust. The
ENG THST LOW PFL is displayed for failed
open/missing nozzle events.
A failed closed nozzle results in normal thrust below
AB and stalls when AB is attempted.
Low or insufficient thrust can be caused by a failed
open, damaged, or missing nozzle or a nozzle control
system malfunction. If thrust is too low to sustain
level flight, turn immediately toward the nearest
suitable runway and establish 250 knots. With a
missing nozzle, level flight may not be possible above
8000 feet MSL.
Thrust available should increase as altitude de
creases. The airspeed at which thrust required forlevel flight is the lowest is approximately 250 knots.
Indications of a nozzle loss are as follows:

S An initial loud bang or pop, similar to a compressor

stall, but rpm is stable above 60 percent; in PRIMIL, engine rpm is approximately 5 percent lower
than normal and FTIT is approximately 250_C
lower than normal; fuel flow is lower than normal;
the nozzle is likely indicated in the full closed
position; and thrust is decreased. Malfunctions of
the exhaust nozzle control system may havesymptoms similar to a missing nozzle, but thenozzle may indicate full open since the nozzle
actuation system is intact.

S Presence of the ENG THST LOW PFL indicatesthat the DEEC has detected the malfunction andhas activated logic to increase the thrust available
in PRI. AB is inhibited. Remain in PRI if possible,
as it should provide a sufficient level of thrust while
also maintaining safe engine operation.

Change 1367

T.O. GR1F16CJ1

S If level flight cannot be attained by 1000 feet above

minimum safe ejection altitude or minimum safe
altitude with the ENG CONT switch in PRI, selectSEC. Set the throttle as required to maintain 250
knots. Continuous operation above 850_C in SEC islikely to result in catastrophic engine failure andfire in as little as 5 minutes.

If thrust is low and a failed open, damaged, or missing
nozzle is suspected:
1.

Throttle-MIL or below.

2.

Stores - Jettison (if required).

3.

Airspeed - 250 knots.

If thrust is sufficient to reach a suitable landing
field:
4.

Land as soon as possible. Plan a flameout
landing. Refer to FLAMEOUT LANDING, thissection.

If unable to reach a suitable landing field and level
flight cannot be maintained by 1000 feet above
minimum recommended ejection altitude or minimum safe altitude, whichever is appropriate:5.

ENG CONT switch - SEC.
NOTE
SEC should only be selected when it
becomes apparent that sufficientthrust cannot be achieved in PRI. SEC
eliminates the additional thrust andthe engine protection benefits provided by the DEEC in PRI. The nozzle
loss logic holds the engine in PRI forthese reasons.

6.

Throttle-As required to maintain 250 knots inlevel flight above minimum recommended
ejection altitude or minimum safe altitude,
whichever is appropriate.

With nozzle loss, catastrophic engine
failure and fire are probable withprolonged high power settings above
850_C FTIT while in SEC.

368

If airspeed drops below 250 knots, trade
altitude to reacquire 250 knots. Do notdescend below minimum recommended
ejection altitude or minimum safe
altitude, whichever is appropriate.
7.

Land as soon as possible. Plan a flameout
landing. Refer to FLAMEOUT LANDING, thissection.

AB Blowout/Failure To Light PW 229
An AB blowout is indicated by the nozzle opening
then closing after the throttle is advanced to AB. If an
AB blowout occurs and the throttle is left in AB, theDEEC automatically recycles the AB up to three
additional times (each cycle indicated by the nozzle
opening and closing). An AB nolight is indicated by
the nozzle failing to start open within 5 seconds ofadvancing the throttle to AB (nozzle remains closed
or shows minimal movement). If an AB no light occurs
and the throttle is left in AB, the DEEC automaticallyattempts to relight the AB up to 3 times. The initialattempt and 3 subsequent no lights could take up to
20 seconds. A combination of no lights and blowout
recycles could take longer. If further AB attempts are
required and the DEEC has completed all recycle
attempts, then the throttle must be retarded to MIL
or below and advanced to AB.
If the AB blowout/failure to light was notaccompanied by an ENGINE FAULT caution light,
flight may be continued. If an ENGINE FAULT
caution light also occurred, refer to ENGINE FAULT
CAUTION LIGHT PW 229 , this section.
ENGINE STALLS PW 229
The three primary causes of a stall are inlet flow
distortion, AB instabilities, and hardware malfunc
tions. During normal aircraft operation, inlet flow
distortion severe enough to cause an engine stall is
not expected. However, under some departure
conditions, inlet flow distortion may induce engine
stalls. Hardwareassociated stalls may result from afailed nozzle, control system malfunctions, antiice
system failed on, or FOD.
Stalls may be caused by an antiice valve failed in the
open position at high thrust settings (throttle above
midrange). The engine should be operable with thiscondition by limiting throttle position to midrange or
below. If flight conditions permit, place the ANTI ICE
switch to OFF.

T.O. GR1F16CJ1
The first indication of an engine stall at high thrust
settings may be a loud bang or pop. At lower thrust
settings, the first indication may be loss of thrust,lack of throttle response, or decreasing engine rpm.
When a stall is sensed, the DEEC cancels the AB (ifthrottle is in AB range), opens the nozzle, anddecreases fuel flow until the stall clears. FTIT andNOZ POS may fluctuate in response to the stallrecovery signal. If the engine auto transfers to SEC,
automatic stall recovery and overtemp protection are
not available. A malfunction such as engine internal
damage or primary control system failure could resultin a stall, an automatic SEC transfer, and possible
FTIT overtemp. Throttle reduction is appropriate asa first response to clear any engine stall.
If the engine stalls at low altitude, an immediateclimb should be initiated, and stores jettison should
be considered. Retarding the throttle may clear thestall. During a high thrust stall that is self recovering,
there will be an immediate thrust loss. In PRI, theDEEC gradually restores thrust to the original level.
If engine response at low altitude is not sufficient tomaintain or gain altitude and a suitable landing field
is not available, ejection may be required.
If a stall occurs at MIL or below, retarding the throttle
may clear the stall. Further throttle movement
should be limited to midrange or below.

engine response. However, nonAB stalls can also be
severe. They may be characterized by bangs, pops,
low intensity or severe engine rumble or vibration. A
momentary nozzle swing to near full open may occur,causing a temporary reduction in thrust, as the
DEEC activates stall recovery. An erratic orangeyel
low flame from the engine exhaust may be present.
This exhaust flame should not be mistaken for anengine fire. If the stall is confirmed, the throttle
should be immediately retarded to IDLE which may
clear the stall. Further throttle movement should be
limited to midrange or below.

Prolonged engine operation with FTIT
in excess of 1000_C can result insignificant engine damage and may
cause a nonrecoverable engine failure.
Engine Stall Recovery PW 229
If an AB stall(s) occurs:
1.

If AB stalls do not clear or stall(s) occurs below AB:
NOTE

ABAssociated Engine Stalls PW 229
ABassociated stalls are normally accompanied by aloud bang or pop and a series of fireballs from theengine exhaust and occasionally the engine inlet. Thisis followed by an erratic flame from the engine exhaust
if the stall is nonrecoverable. These characteristics
could be mistaken for an aircraft fire. Whenever a stall
occurs while operating in AB, the DEEC automaticallycancels AB and activates stall recovery. This may be
accompanied by a nozzle swing to full open for a few
seconds and an associated temporary reduction ofthrust. The throttle should be snapped out of AB to
MIL. This action usually clears the stall and restores
normal operation; however, stalls may continue at MIL
and can be severe. They may be characterized by bangsor pops of low intensity or engine vibrations severe
enough to preclude reading engine instruments. Refer
to NONAB ENGINE STALLS PW 229 , this section.

Throttle-Snap to MIL.

NonAB stalls may be inaudible.
2.

Throttle-IDLE.

3.

ANTI ICE switch-OFF when conditions
permit.
NOTE
Stalls may be caused by antiice valve
failing to close at high thrust setting(throttle above midrange).

If stalls continue at idle and engine rpm is less than 60
percent with no rpm response to throttle movement:
4.

Throttle-OFF. Initiate airstart. Refer to AIRSTART PROCEDURES PW 229 , this section.

NonAB Engine Stalls PW 229
NonAB stalls may occur if the engine is malfunctioning, particularly during throttle transients near
IDLE. NonAB stalls are often a symptom of a serious
engine problem. NonAB stalls may be inaudible; thefirst indication may be a lack of throttle response
which may be difficult to differentiate from abnormal

Shutting down the engine with anengine alternator failure (indicated byzero or erroneously low rpm, illuminated
SEC caution light, illuminated ENGINE
warning light, and normal thrust)
results in no ignition for an airstart.
369

T.O. GR1F16CJ1
If nonAB stall(s) clears:
5.

Throttle-Midrange or below.
If a nonAB stall clears, maintain throttle at
midrange or below unless required to sustain
flight.

6.

Land as soon as possible.

If AB stall(s) clears:
2.

Throttle-As required.
If an AB stall clears, the engine is safe tooperate in the IDLE to MIL range, provided
no other abnormal indication is observed.
Attempt further AB operation only if needed
to sustain flight.

INLET BUZZ PW 229
Inlet buzz occurs at supersonic airspeeds if the engine
control system fails to maintain adequate engine rpm
when the throttle is retarded below MIL. Inlet buzz
causes moderate to severe vibration within the
cockpit and probably results in multiple engine stalls.
If inlet buzz occurs, do not move the throttle until
subsonic. Decrease airspeed to subsonic as quickly as
possible by opening the speedbrakes and increasing g.
If engine stalls occur and persist, the throttle should
be retarded to IDLE when subsonic. If the stalls do not
clear, the engine must be shut down and restarted.
ENGINE FAILURE OR FLAMEOUTPW 229Engine failures can result in rpm decrease with noabnormal vibration or sound (flameout), rpm
decrease with abnormal vibration and/or stalls, or
stable rpm with abnormal vibration and/or low
thrust.
If the engine flames out, fuel starvation ormechanical failure has occurred. A flameout isindicated by decrease in FTIT and engine rpm
decaying below approximately 60 percent. Loss ofthrust and lack of response to throttle movement
confirm the flameout. The ENGINE warning lightilluminates when engine rpm is below 55 percent.
Additionally, the MAIN GEN and STBY GEN lights
illuminate below 45 percent rpm and the EPU should
start running. Do not mistake a loss of ECS noise as
an engine flameout.
If the reservoir tanks do not contain fuel, an airstartis impossible. If fuel starvation was due to atemporary lack of fuel, restart should be possible. If
fuel quantities appear normal, the flameout may
have been caused by fuel contamination. In this case,placing the throttle to OFF may clear thecontaminated fuel and allow an airstart.
370

Main fuel pump failure or tower shaft geartrain
failure also causes flameout. Both present similar
symptoms:an abrupt decrease of indicated fuel flow
to less than 500 pph; loss of main generator, standby
generator, and FLCS PMG and EPU activation; nothrottle response; and illumination of the SEC
caution light even though the ENG CONT switch is in
C DF PRI, DR NORM.
If only the main fuel pump has failed, the rpm
indication reflects a gradual spooldown. The JFS canbe started and the engine can be motored at
approximately 25 percent rpm. If the SEC caution
light remains on (with ENG CONT switch in C DF
PRI, DR NORM and engine rpm at 12 percent orabove), the engine probably cannot be restarted;
therefore, place primary emphasis on a flameout
landing while continuing airstart attempts. If unable
to make a flameout landing, refer to EJECTION, thissection.
Tower Shaft Failure PW 229
Failure of the engine tower shaft or its associatedgeartrain results in loss of all rotation to the engine
gearbox and the ADG. Loss of rotation to the engine
gearbox renders the engine alternator, main fuel pump,and oil pump inoperative resulting in a zero rpm
indication, zero oil pressure, illumination of the
ENGINE warning and SEC caution lights, and engine
flameout due to fuel starvation. The initial symptoms
are similar to main fuel pump failure; however, theprimary difference is that the rpm and oil pressure
indications drop immediately to zero with a tower shaft
failure since the engine alternator is not being driven.
Additional symptoms caused by loss of rotation to the
ADG include loss of hydraulic systems A and B, mainand standby generators, and FLCS PMG andsubsequent activation of the EPU. It may be possible toregain engine operation using the JFS and performing
an SEC airstart. The JFS drives the ADG and theengine gearbox (through the PTO shaft), restoringrotation to both hydraulic pumps, FLCS PMG (at areduced output), main fuel pump (SEC caution lightgoes off in PRI until SEC is selected), engine alternator(cockpit rpm signal, DEEC power, and engine ignition),and oil pump (oil pressure increases). Without the loadof the engine, the JFS produces an rpm indication
fluctuating between 3050 percent which is the speed ofthe engine alternator, not the actual engine rpm. Thisrpm may be high enough to restore standby generator
power; however, main generator power may cycle on
and off line with the rpm fluctuations. If the ENG
CONT switch is still in C DF PRI, DR NORM, the SEC
caution light goes off when fuel pump pressure isrestored; however, a PRI airstart is not possible since

T.O. GR1F16CJ1
the rpm signal to the DEEC is in error. Perform an
SEC airstart. Since the JFS is not preserving rpm,
maintain 250 knots minimum during the airstart
attempt, which should assure adequate actual engine
rpm for the airstart.
Low Altitude Engine Failure or Flameout PW 229
Refer to figures 37 and 38. Initial reaction to any
malfunction at low altitude should be to trade excess
airspeed for altitude. Higher altitude translates
directly to either additional time to achieve anairstart or to additional glide range to reach a suitablelanding field. At low airspeed, the climb may be only
enough to insure a safe ejection altitude. Above 350
knots, more time is available by a zoom climb using
a 3g pullup to 30degree climb approaching thedesired airspeed (use approximately 50 knots lead
point) and then initiating a zero g pushover. Below
350 knots and above the minimum recommended
ejection altitude, more time is available by perform
ing a constant altitude deceleration to the desired
airspeed. If below the minimum recommended
ejection altitude and below 350 knots, primaryconcern should be to trade excess airspeed for altitudein preparation for ejection.
If required, jettison stores as soon as possible to aid ingaining or maintaining altitude and maneuver
toward a suitable landing field, if available. If thezoom results in an altitude below 4000 feet AGL,
there will probably be insufficient time to achieve anairstart prior to minimum recommended ejection
altitude. In that case, primary consideration should
be given to preparing for ejection; do not delayejection below 2000 feet AGL. If the zoom results inan altitude between 400010,000 feet AGL, there is
probably time for one airstart attempt prior tominimum recommended ejection altitude. This
attempt shall be performed in the control mode
selected by the DEEC.
If low altitude engine failure or flameout occurs:
1.

Zoom.

2.

Stores-Jettison (if required).
If stores jettison is attempted after main andstandby generators drop off line but before
EPU generator powers the SMS (approxi
mately 5 seconds delay), stores will notjettison.
NOTE
Visually confirm the stores have jetti
soned and jettison again if required.

3.

Perform airstart (if altitude permits). Refer to
AIRSTART PROCEDURES PW 229 , this section.

Below 4000 feet AGL, there may beinsufficient time to perform an airstartprior to minimum recommended ejec
tion altitude.
AIRSTARTS PW 229
Refer to figure 39. Factors such as altitude, airspeed,
weather, etc., must be considered in determining
whether to try an airstart, accomplish a flameout
landing, or eject. Jettisoning of stores reduces
altitude loss during an airstart and improves glide
ratio during a flameout landing.
Oil pressure is directly related to rpm. Do not confuse
a low oil pressure indication due to windmilling rpm
as an oil system malfunction.
If the engine seized due to an oil system malfunction
or flamed out due to total fuel starvation ormechanical failure, either a flameout landing orejection is required.
The most likely reason to perform an airstart is thatthe engine has shut down due to a PRI system failureor hardware failure or to clear a stall. The DEEC
assesses any faults or internal failures andautomatically transfers to SEC, if required. The first
airstart attempt should be made in the engine control
mode selected by the DEEC except when a tower shaft
failure is suspected it shall be performed in SEC.
Procedures for SEC and PRI airstarts are identical
except for ENG CONT switch position and JFS assist
minimum airspeed requirements.
There are two airstart options available. One option
is a spooldown airstart, for which the throttle isadvanced from OFF to midrange as rpm isdecreasing. There is no specific envelope within
which a spooldown airstart should be initiated, but
once the decision to perform an airstart is made,airspeed should be adjusted to maintain 250400
knots/0.9 mach. The secondary option is a
JFSassisted airstart which differs from a spool
down airstart in that once the JFS RUN light is on
and PRI mode is confirmed, airspeed can be reduced
to achieve maximum range or maximum endurance( C 200 or 170, D 205 or 175 knots respectively, plus
5 knots per 1000 pounds of fuel/store weights andplus 5 knots if CFT's are installed). The minimum
airspeed for PRI JFSassisted airstarts is 170
knots. The minimum airspeed for all SEC airstarts,
including JFS assisted, is 250 knots.
Change 1371

DATA BASIS ESTIMATEDENGINE F100-PW-229

CONFIGURATION:

CONDITIONS:GW = 23,000-25,000 LB
DI = 0-50
LG UP

WINDMILLING OR SEIZED ENGINE
30-DEGREE CLIMB MAINTAINED TO 170/250 KIAS

12
3G PULLUP TO 30
ZOOM CLIMB TO
ACHIEVE:
10250 KIASINITIAL ALTITUDE

1000 FT AGL

170 KIAS

8

6

4

2

0

0

160

180

200

220

240

260

280

300

320

340

INITIAL AIRSPEED

360

KIAS

Figure 37.

380

400

420

440

460

480

500
1F-16X-1-4028X

T.O. GR1F16CJ1

372

Low Altitude Zoom Capability

Low Altitude Airstart Capability
DATA BASIS ESTIMATEDENGINE F100-PW-229

CONFIGURATION:

CONDITIONS:GW = 23,000-25,000 LB

30 DIVE TO DESCENT KIAS OR 3G PULLUP
TO 30 ZOOM CLIMB INITIATED FROM THE
AIRSPEED/ALTITUDE EXISTING AT FIRSTRECOGNITION OF ENGINE FAILURE (60 PERCENT RPM)

DI = 0-50LG

UP

60 SECONDS ASSUMED AFTERTHROTTLE ADVANCE TO ACHIEVE
USABLE THRUSTDESCENT AIRSPEED IS 250 KIAS(JFS RUN LIGHT ON)

AIRSTART INITIATED AT START OF DIVE OR
ZOOM
10
30

9

DIVE TO 250 KIAS

3G PULLUP TO 30

ZOOM CLIMB

260

300

INITIAL ALTITUDE

1000 FT AGL

8
7
6
5
4
3
2

373

0

MINIMUM RECOMMENDED
EJECTION ALTITUDE

0

100

120

140

160

180

200

220

240

280

320

INITIAL AIRSPEED

340

360

380

400

420

440

460

480

500

520

KIAS
GR1F16CJ11120X37

Figure 38.

T.O. GR1F16CJ1

1

T.O. GR1F16CJ1Engine RPM and FTIT Response During
Spooldown and Airstart
ENGINE F100-PW-229

CONDITIONS:DRAG INDEX = 0
KIAS = 250

GW = 20,000 LB
NO WIND

A

50

30 DIVE
POINT B TO
SL = 7.9 MIN/40 NM

ALTITUDE

1000 FEET40

PRI SPOOLDOWN OR
SEC SPOOLDOWN/JFS
ASSIST-250 KIAS

30BMIN KIAS WITH JFSRUN LIGHT ON
AND PRI MODE
AIRSTART ENVELOPE
CONFIRMED-170

400 KIAS/0.9 MACH

20JFS ENVELOPE
10

0

0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

MACH NUMBER
OPTIMUM FLIGHT PATH DURING AN AIRSTART (TYPICAL)Engine out descent flight path maintains the aircraft in the required airstart envelope.
A 30-degree dive to 400 KIAS/0.9 mach is used to quickly reduce altitude to below
30,000 feet where airspeed can be reduced to 250 KIAS (Point A to B).

1F-16X-1-1039A

Figure 39.(Sheet 1)374

T.O. GR1F16CJ1Engine RPM and FTIT Response During
Spooldown and Airstart
ENGINE F100-PW-229

1000
250/350 KIAS
10,000 FEET

C

800

FTIT

600
400

250/350 KIAS
35,000 FEET

200
100
350 KIAS
10,000 FEET

%

60

RPM

80

40

250 KIAS
35,000 FEET
350 KIAS
35,000 FEET
25%

20
0

250 KIAS
10,000 FEET
0

20
40
TIME
SECONDS

60

RATE OF ENGINE RPM AND FTIT DECAY
AS A FUNCTION OF ALTITUDE AND
AIRSPEEDAt low altitude, regardless of airspeed,
the spooldown rate is rapid. Spooldown
rate is slower at higher altitudes asairspeed is increased.

1F-16X-1-1040X

Figure 39.(Sheet 2)

375

T.O. GR1F16CJ1Engine RPM and FTIT Response During
Spooldown and Airstart
ENGINE F100-PW-229
THROTTLE TO OFF

1200

THROTTLE TO MIDRANGE AT 25%

FTIT

C

1000

LIGHTOFF

800
600
400
200

100

RPM

%

80
60
40
25%
20
0
0

10

20

30

40

TIME

SECONDS

50

60

70

NORMAL 25 PERCENT SPOOLDOWN AIRSTART ENGINE RPM AND FTIT
TIME TRACES FOR 250 KIAS/20,000 FEET
Light-off normally occurs within 5 seconds after the throttle is advanced to midrange.
However, engine rpm and FTIT turnaround are slow, making light-off subtle or difficult
to detect. Engine rpm stabilizes momentarily after light-off and FTIT may increase,stabilize, or decrease as the engine rpm increases.
1F-16X-1-1041X

Figure 39.(Sheet 3)376

T.O. GR1F16CJ1Engine RPM and FTIT Response During
Spooldown and Airstart
ENGINE F100-PW-229
THROTTLE TO OFF

1200

THROTTLE TO MIDRANGE AT 50%
1000

FTIT

C

LIGHTOFF
800
600
400
200

100

RPM

%

80
60

50%

40
20
0
0

5

10

15
TIME

20

25

30

SECONDSNORMAL 50 PERCENT SPOOLDOWN AIRSTART ENGINE RPM AND FTIT
TIME TRACES FOR 250 KIAS/20,000 FEET
Light-off normally occurs within 5 seconds after the throttle is advanced to midrange.
However, engine rpm and FTIT turnaround are slow, making light-off subtle or difficult
to detect. Engine rpm stabilizes momentarily after light-off and FTIT may increase,stabilize, or decrease as the engine rpm increases.
1F-16X-1-1042X

Figure 39.(Sheet 4)

377

T.O. GR1F16CJ1NOTEFIn most cases, the JFS will engage andbegin to arrest rpm spooldown rate at
approximately 30 percent. Engine rpm
should stabilize on the JFS at aminimum of 22 percent.

FPRI operation is confirmed by SEC

caution light not illuminated.

There are critical requirements which apply to any
airstart attempt. The most important is engine rpm.
High engine rpm provides the best chance of a
successful restart. Therefore, do not delay the
initiation of an airstart in an attempt to reach aparticular flight condition. Initiate an airstart assoon as it becomes apparent that engine rpm has
decayed below inflight idle (approximately 60
percent rpm). Illumination of the ENGINE warning
light, engine instrument indications, and no
response to throttle movement can help confirm aflameout. The best conditions for an airstart are at
30,000 feet MSL or below, 250 knots or more, andhigh rpm. In general, the success of an airstart isincreased with higher engine rpm, lower altitude,
and higher airspeed.
The throttle should always be advanced to midrange
before 25 percent rpm regardless of FTIT, altitude,
airspeed, JFS assistance, or engine control mode toprevent rpm decreasing below 12 percent. If rpm
decreases below 12 percent, fuel flow from the main
fuel pump and ignition power from the engine
alternator are lost allowing a further decrease in rpm.
As much as 350 knots or more is required to prevent
rpm from decreasing below 12 percent. If rpm isallowed to decrease to near zero, 400 knots or more
may be required to regain 12 percent. This requires a
great amount of altitude which may not be available.
If rpm decreases to zero, a seized rotor may occur, due
to temporary engine thermal conditions, after which
it will not rotate even with high airspeeds or byengaging the JFS. In general, rpm decay rate can be
decreased by increasing airspeed; however, below
30,000 feet MSL, higher airspeeds do not significantly affect the rpm decay. Therefore, maintain 250
knots below 30,000 feet MSL for spooldown airstarts.
This airspeed does not maintain rpm above 12
percent; however, it does the best tradeoff between
the rate of rpm spooldown and loss of altitude.
Stabilizing or increasing rpm is normally the firstindication of an engine lightoff. Lightoff normallyoccurs within 5 seconds after advancing the throttle.

378

However, rpm turnaround is slow, making lightoff
subtle and difficult to detect. FTIT response is slow
and may reverse trend one or more times during thestart sequence. Neither condition should be confusedwith a hung start.
DEEC overtemperature protection logic attempts tolimit FTIT to 870_C, which may result in decreasing,hung, or slowly increasing engine rpm. If a hung startoccurs (stabilized FTIT 870_C or less, rpm stabilized
below 60 percent), increase airspeed to a maximum of
400 knots/0.9 mach if altitude allows. If hung start
continues or there is no throttle response, reinitiate
the airstart with the ENG CONT switch in SEC when
below 30,000 feet MSL.
During an SEC start, 4560 seconds is required forengine lightoff and acceleration to midrange fromthe time the throttle is advanced from OFF.
If a SEC airstart is initiated at high RPM (above 40
percent), as much as 30 seconds following throttle
advancement may be required for RPM to stabilize or
begin to increase, indicating a successful light. This
slow engine response should not be confused with anolight response in which case RPM would continue
to decrease.
SEC airstarts initiated at lower RPM show positive
RPM turnaround within 15 seconds of placing
throttle to midrange. RPM continuing to decrease
after 15 seconds from throttle advancement or
stabilizing at a minimum of 22 percent with the JFS
RUN light on, indicates a nolight has occurred.
High Altitude Airstart Considerations PW 229
Refer to figure 39. At 30,000 feet MSL and above, theairstart (PRI or SEC) should be initiated by retarding
the throttle to OFF then advancing to midrange assoon as possible regardless of FTIT or airspeed.
Always advance the throttle to midrange by 25
percent engine rpm. Dive to obtain 400 knots/0.9
mach to minimize rpm spooldown rate and quickly
decrease altitude to less than 30,000 feet MSL. If
lightoff indications are not noted within 20 seconds
in PRI or 30 seconds in SEC after advancing thethrottle, or if FTIT exceeds 870_C, retard the throttle
to OFF and reinitiate the airstart with the ENG
CONT switch in C DF PRI, DR NORM, regardless ofwhich control mode the engine is presently in. If ahung start occurs (rpm stable with FTIT stable at
870_C or less), keep the throttle at midrange until
below 30,000 MSL then reinitiate the airstart with
the ENG CONT switch in SEC.

T.O. GR1F16CJ1
At high altitudes, the dive should be at approximately
30 degrees to gain or maintain 250 knots below 30,000
feet. Once established, approximately 510 degrees ofdive should maintain airspeed. Note that airspeed
should not be reduced to less than 250 knots until the
JFS RUN light is on, and PRI mode is confirmed.
Unless an airstart is obviously impossible (total lackof fuel, engine seizure, etc.), do not become tempted toestablish a maximum range or maximum endurance
glide. The first consideration should be an immediatespooldown airstart attempt even if the engine failed
for no apparent reason. If airstart airspeed is not
maintained, rpm decreases at a faster rate. The only
airstart option available is then a JFSassisted
airstart. Time constraints due to EPU fuel consumption must be considered. A maximum range or
maximum endurance glide from above approximately
35,000 feet may exhaust EPU fuel prior to landing.
(Refer to T.O. GR1F16CJ11, figure A63.)

Following a zoom climb, plan to arrive at 250 knots.
Airspeed may be reduced to achieve maximum
range or maximum endurance (200 or 170 knots,
respectively) only if PRI mode is confirmed and the
JFS RUN light is on. If SEC mode is confirmed or
tower shaft failure is suspected, maintain 250 knots
minimum. If a higher airspeed is maintained or an
attempt is made to gain airspeed to delay the rpm
decay, available time may be reduced to the pointthat an airstart is not possible. During any low
altitude airstart attempt, constantly evaluate
altitude above the ground relative to airstartsuccess. Do not delay ejection below 2000 feet AGL
unless the engine is producing thrust capable ofmaintaining level flight or safely controlling thesink rate or unless a flameout landing can beaccomplished.

The reason for most airstart failures above 30,000
feet MSL is the engine inability to light off and not
due to the DEEC's inability to control the start.
Therefore, all second airstart attempts above 30,000
feet MSL should be made with the ENG CONT switch
in C DF PRI, DR NORM. The engine does not have alightoff problem below 30,000 feet; therefore, allsecond airstart attempts below 30,000 feet MSL
should be made with the ENG CONT switch in SEC.

To perform an airstart, retard throttle to OFF, thenadvance throttle to midrange.

When below 20,000 feet MSL, select JFS START 2.
Activating the JFS above 20,000 feet is prohibitedsince successful JFS start/motoring of engine isunlikely and the brake/JFS accumulators will be
depleted. If the JFS RUN light is on and PRI mode isconfirmed, airspeed may be reduced to achieve
maximum range or maximum endurance. If the JFS
RUN light is on and SEC mode is confirmed,
maintain 250 knots minimum. With the JFS
running, EPU fuel consumption is also reduced.
Low Altitude Airstart Considerations PW 229
Due to the limited time available and the rapid rpm
spooldown rate at low altitude, some additional
considerations are required. Below approximately
10,000 feet MSL, rpm decreases rapidly regardlessof airspeed and remains between 5025 percent for
only 510 seconds; therefore, rpm should be closely
monitored. Advance the throttle to initiate theairstart before rpm goes below 25 percent
regardless of FTIT indication. This action should
insure that lightoff occurs prior to 12 percent rpm.
Start the JFS immediately after advancing thethrottle (if airspeed is below 400 knots).

Airstart Procedures PW 229

NOTE
If the throttle is retarded to OFF to clear
a stall, it should be maintained in OFF fora few seconds to allow the stall to clear.
Start the JFS below 20,000 feet MSL and below 400
knots immediately after the throttle is advanced to
midrange to initiate the spooldown airstart.
If the JFS stops running or fails to run within 30
seconds, do not reattempt a JFS start until thebrake/JFS accumulators have time to recharge. Allow
1 minute of engine rotation (either windmilling or
JFS assisted) at 12 percent rpm or above to insurethat the brake/JFS accumulators are fully recharged.
Recharging begins 34 seconds before the JFS RUN
light illuminates or 30 seconds after selecting a start
position (in the event of a JFS failure to run).
Recharging begins regardless of JFS switch position.
In the event of a JFS shutdown, the JFS switch does
not relatch in either start position while the JFS isspooling down. Spooldown from full governed speed
takes approximately 17 seconds. The JFS switch
must be cycled to OFF and then to START 2 toreinitiate a JFS start. It is possible to complete thespooldown before the brake/JFS accumulators arerecharged if the JFS ran for only a short time.
When the airstart is completed, turn the JFS off (iftower shaft failure is not suspected). Reset the main
generator using the ELEC CAUTION RESET buttonand verify MAIN GEN and STBY GEN lights are off.
Cycle the EPU switch to OFF, then back to NORM.

379

T.O. GR1F16CJ1
NOTE

With engine failure or flameout,
OBOGS is inoperative. Activate EOS if
OXY LOW warning light illuminates
above 10,000 feet cockpit altitude.

Visually confirm the stores havejettisoned and jettison again if required.
If a no light, hot start, or stall occurs:

To accomplish an airstart:

5.

Throttle-OFF.

1.

6.

ENG CONT switch-SEC if below 30,000 feet
MSL (250 knots minimum).

Throttle-OFF, then midrange.
NOTE

NOTE

FTIT will decrease rapidly whenthrottle is OFF.
2.

Airspeed-As required.
Above 30,000 feet MSL, dive at 400
knots/0.9 mach. Below 30,000 feet MSL,
establish approximately 250 knots. When
below 20,000 feet MSL with the JFS RUN
light on and PRI mode confirmed, airspeed
can be reduced to achieve maximum range
or maximum endurance ( C 200 or 170, D
205 or 175 knots, respectively, plus 5 knots
per 1000 pounds of fuel/store weights andplus 5 knots if CFT's are installed).
NOTE
If maximum gliding range is not a factor,
consider maintaining 250 knots above
10,000 feet AGL to reduce rpm spooldown
rate (in case of JFS failure). Below 10,000
feet AGL with the JFS RUN light on and
PRI mode confirmed, maintain maximum
range or maximum endurance airspeed.

3.

JFS switch-START 2 below 20,000 feet MSL
and below 400 knots.

FPlace the ENG CONT switch to SEC

prior to placing the throttle tomidrange, otherwise a start anomaly
may result.

FThe proximity of the ENG CONT

switch to the JFS switch makes the
JFS switch susceptible to being
bumped to OFF when selecting SEC.

7.

If a hung start occurs:
8.

Airspeed-Increase (maximum of 400 knots/
0.9 mach).

If hung start continues or there is no throttle
response:9.

Throttle-OFF when below 30,000 feet MSL.

10. ENG CONT switch-SEC (250 knots mini
mum).

NOTE

NOTE

FIf the JFS switch is erroneously placed

FPlace the ENG CONT switch to SEC

prior to placing the throttle tomidrange, otherwise a start anomaly
may result.

to START 1, leave it there.

FIf the JFS RUN light does not illuminate

or goes off once illuminated, place the
JFS switch to OFF and reattempt
START 2 when the brake/JFS accumu
lators are recharged. The JFS switch
does not relatch in either start position
while the JFS is spooling down.

4.

Throttle-Midrange.

Stores-Jettison (if required).
If stores jettison is attempted after main
generator drops off line but before EPU
generator powers the SMS (approximately 5
seconds delay), stores will not jettison.

380Change 1

FThe proximity of the ENG CONT

switch to the JFS switch makes the
JFS switch susceptible to being
bumped to OFF when selecting SEC.

11.

Throttle-Midrange.

If engine does not respond normally after airstart iscompleted:
12. Refer to FLAMEOUT LANDING, this section.

T.O. GR1F16CJ1
If engine responds normally:

Do not turn JFS or EPU off if indicated
rpm is below 60 percent with adequate
thrust (e.g., tower shaft failure).
12. JFS switch-OFF.
13. ELEC CAUTION RESET button-Depress.
Verify MAIN GEN and STBY GEN lights areoff.
14. EPU switch-OFF, then NORM.
15. ADI-Check for presence of OFF and/or AUX
warning flags.
If warning flag(s) is in view, refer to EGIFAILURE, this section.
16. Throttle-As required.
NOTE
If the SEC caution light is on, refer to
SEC CAUTION LIGHT, this section.
17. Land as soon as possible.
18. Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.
ENGINE MALFUNCTIONS GE 129
The EMS compares expected versus actual engine
operation. The purpose of the EMS MFL is to provide
maintenance personnel with an early indication ofan engine condition which requires correction. No
action is required for an engine MFL at anytime
during a flight.
Low altitude, for engine malfunction purposes, isgenerally defined as 10,000 feet AGL or below. If an
engine malfunction is suspected, the initial reaction
should be to trade excess airspeed for altitude. Unlessa suitable airfield is within gliding distance, turns
should be avoided as they decrease the amount oftime/altitude available to successfully recover engine
performance or prepare for ejection. Optimizing theexchange of airspeed for altitude must be a priority
action for any engine malfunction. Above 310 knots,
more time is available by performing a zoom climb
using a 3g pullup to 30degree climb until
approaching the desired airspeed (use approximately
50 knots lead point) and then initiating a zerogpushover. Below 310 knots and above the minimum

recommended ejection altitude, more time isavailable by performing a constant altitude deceleration to the desired airspeed. If below the minimum
recommended ejection altitude and below 310 knots,
primary concern should be to trade excess airspeed
for altitude in preparation for ejection. If appropriate,
jettison stores as soon as possible.
For any situation where automatic activation of the
EPU is relied upon, verify that the EPU run light is
on to insure that the EPU has started. If the EPU run
light is off, position the EPU switch to ON.
SEC is a highly reliable backup engine operating
mode. However, in certain circumstances the engine
may malfunction in SEC following an auto transfer,or continue to malfunction in SEC following a manualtransfer. Returning to PRI may result in proper
engine response, due to either the resetting ofinternal hydromechanical control pressures or to theresetting of digital engine control logic.
Idle thrust in SEC during ground operation isapproximately twice that in PRI. After landing in
SEC, consider minimizing taxi distance and consider
following HOT BRAKES procedures, this section.
Engine Fire GE 129
Generally, the first indication of fire in the engine
compartment is the ENG FIRE warning light.
Abnormal fuel indications (quantity/flow) may also
be present. FTIT probably will not be higher thannormal. Explosions, vibrations, or engine instrument
fluctuations are usually indicative of a serious engine
problem; engine failure may be imminent. Immediate
action should be taken to reduce thrust to theminimum practical level after attaining safe ejection
parameters. If within gliding distance of a suitablerunway, consider shutting the engine down. Suffi
cient time should exist to analyze the situation andmake an ejection versus land decision. The ejection
decision should be based on visual and/or cockpit
indications that the fire is persisting. Cockpit
indications include continued illumination of the
ENG FIRE warning light and subsequent FLCS
malfunctions/degraded flight controls or subsequentloss of either hydraulic system.
Fires can also occur in the nozzle area when using AB.
These fires are the result of portions of the nozzle
failing which allows the AB plume to burn throughthe nozzle. Ventilation should inhibit forward
movement of the fire into and through the engine bay.
Since these fires are aft of the detection circuit, the
ENG FIRE warning light will not illuminate.
Additionally, the nozzle position indications arenormal, and there are no vibrations or instrument
fluctuations. In most cases, these ABrelated nozzle
Change 1381

T.O. GR1F16CJ1
fires are detected by someone outside the aircraft(wingman, tower, etc.). When operating in AB and afire is reported at the rear of the aircraft, retard
throttle below AB immediately. This action should
extinguish a nozzle fire within approximately 3045
seconds and minimize damage to the aircraft skin,speedbrakes, nozzle, and flight controls; however,nozzle damage may result in a noticeable thrust loss.

2.

If OVERHEAT caution light remains on (or detectcircuit checks bad) and EPU is running:3.

If on takeoff and the conditions permit:
1.

Abort.

If takeoff is continued:
1.

2.

Climb.
Maintain takeoff thrust until minimum
recommended ejection altitude is attained
and then throttle to minimum practical.

4.

OXYGEN-100%.

5.

AIR SOURCE knob-OFF.
External fuel cannot be transferred in OFF
or RAM. Consider jettisoning tanks todecrease drag if range is critical and the ECS
cannot be turned on for short periods of timeto transfer fuel.

Stores-Jettison (if required).

Throttle-Minimum practical.
If fire occurred in AB, ENG FIRE warning
light may not illuminate. Fire should
extinguish after throttle is retarded; how
ever, nozzle damage may result in lower thannormal thrust.

With the ECS shut down or the AIR
SOURCE knob in OFF or RAM, thegsuit does not inflate and PBG isdisabled.

If ENG FIRE warning light goes off:
4.

FIRE & OHEAT DETECT button-Depress.
Determine if fire detection circuit isfunctional.

If fire persists:
5.

6.

Descend to below 25,000 feet and reduce
airspeed to below 500 knots.

When airspeed is reduced and cockpit is depressu
rized:
7.

Eject.

If fire indications cease:5.

EPU switch-OFF (if feasible).
If the EPU was manually turned on, consider
turning it off to determine if it is the sourceof the overheat condition. If the OVERHEAT
caution light remains on, the EPU should beturned back on.

If OVERHEAT caution light remains on (or detectcircuit checks bad):At a safe altitude:
3.

FIRE & OHEAT DETECT button-Depress.
Determine if fire detection circuit isfunctional.

Land as soon as possible.

AIR SOURCE knob-RAM (below 25,000feet).
External fuel cannot be transferred in OFF or
RAM. Consider jettisoning tanks to decreasedrag if range is critical and the ECS cannot be
turned on for short periods of time to transfer
fuel.

OVERHEAT Caution Light GE 129
Detection of an overheat condition in the engine
compartment, ECS bay, MLG wheel wells, or EPU
bay illuminates the OVERHEAT caution light.
Accomplish as many of the following as required toextinguish the light. If the light goes off, verify the
integrity of the detection circuit by depressing the
FIRE & OHEAT DETECT button and land as soon as
possible.
If OVERHEAT caution light illuminates:
1.

Throttle-Minimum practical.

382Change 1

With the ECS shut down or the AIR
SOURCE knob in OFF or RAM, the gsuit
does not inflate and PBG is disabled.
8.

Nonessential electrical equipment-Off.
NOTE
If in VMC and the ADI and HSI are notrequired for flight, the INS should beconsidered nonessential.

T.O. GR1F16CJ1
If OVERHEAT caution light still remains on (ordetect circuit checks bad):
9.

TANK INERTING switch-TANK INERTING even if Halon is not available.

10. LG handle-DN (300 knots/0.65 mach maximum). (Use DN LOCK REL button if required.)

If LG handle does not lower, selectBRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND.
11.

Land as soon as possible.

Oil System Malfunction GE 129
An oil system malfunction may be indicated by the
EMS, the OIL pressure indicator, the NOZ POS
indicator, and the HYD/OIL PRESS warning light.
An oil system malfunction is characterized by apressure (including fluctuations) below 15 psi at
IDLE or 25 psi at MIL, pressure above 65 psi at any
throttle setting, pressure fluctuations greater than
"5 psi, or continuous presence of the ENG LUBE
LOW PFL.

performing a fixed throttle climb. If the altitude
gained is insufficient for a safe cruise, a slow throttle
advance followed by a slow throttle reduction, when
at a sufficient cruise altitude, may be performed. Do
not make subsequent throttle movements unlessrequired to sustain flight.
Anticipate engine seizure approximately 5 minutes
after oil pressure drops below 10 psi. Use minimal
throttle movement and reduced thrust settings to
maximize time. Minimize maneuvering g to minimize
loads. Plan an approach which allows a flameout
landing from any position should engine seize. Refer
to SIMULATED FLAMEOUT (SFO) LANDING and
FLAMEOUT LANDING, this section.
The EPU should be manually activated if engine
seizure is anticipated (i.e., oil pressure decreases
below 10 psi); otherwise, if the EPU does not start
automatically when the engine seizes, the short timeremaining before loss of control may be inadequate
for recognition of the EPU failure and corrective
action. Monitor hydrazine use after activating the
EPU. If consumption rate is too high, cycle EPU
switch to OFF, then NORM to conserve hydrazine. Be
prepared to place the EPU switch back to ON if theengine seizes.
If an oil system malfunction is suspected:
1.

Engine oil level decreasing below approximately 40
percent of normal capacity for 15 seconds results inactivation of an ENG LUBE LOW PFL. Further oillevel decrease may cause the exhaust nozzle to fail toan aerodynamically balanced position and there
may be a decrease in thrust and a larger than normal
NOZ POS indication. The HYD/OIL PRESS warning
light may not illuminate until most of the usable oil
is lost.

Range-Maximize.
If oil pressure is 10 psi or above and thenearest suitable airfield is not within
gliding distance, consider using MIL or AB
thrust to quickly gain altitude and/or
decrease distance to the nearest suitable
airfield. If oil pressure is below 10 psi, attain
desired cruise conditions using minimum
throttle movement and then slowly retard
to the minimum setting required.

NOTE
FWhen oil pressure is below 10 psi,

An ENG LUBE LOW PFL can occur as
the result of sustained longitudinal
acceleration during takeoff or duringother maneuvering.

throttle movement, high throttle settings, or high airspeeds may accelerate
or cause engine seizure.

FDo not start the JFS if engine seizure

High airspeeds (especially at low altitude), highthrottle settings, and/or throttle movements are
undesirable when oil pressure is abnormally low.
Initial reaction to an oil system malfunction at low
altitude should be to gain altitude and, if oil pressure
is low, reduce airspeed. If oil pressure is 10 psi or
above, consider using MIL or AB thrust to quickly
attain desired cruise altitude. If oil pressure is below
10 psi, increase altitude and reduce airspeed byhas occurred or is anticipated. Start
ing the JFS may result in no brake/JFS
accumulator pressure for the brakes.

2.

Plan to land at the nearest suitable airfield.

If oil pressure is low (with or without an ENG LUBE
LOW PFL):
3.

Stores-Jettison (if required).
383

T.O. GR1F16CJ1
4.

5.

6.

EPU switch-ON, if oil pressure decreases
below 10 psi.
Monitor hydrazine use. If consumption rateis too high, cycle EPU switch to OFF, thenNORM to conserve hydrazine. Be preparedto place EPU switch back to ON if the engine
seizes.
Land as soon as possible.
Plan to fly an SFO. Refer to SIMULATEDFLAMEOUT(SFO)LANDING
andFLAMEOUT LANDING, this section.
Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.

If oil pressure is fluctuating or high (without an ENG
LUBE LOW PFL):
3.

Land as soon as practical at the nearest
suitable airfield.

If oil pressure is normal and ENG LUBE LOW PFL
occurred:3.

FACK, DR FAULT ACK button Depress to acknowledge fault.

4.

Turn toward the nearest suitable airfield.

5.

Maintain straight, level, and unaccelerated
flight for 15 seconds.

6.

C

DF

C DF FACK, DR FAULT ACK button Depress for fault recall.
Presence of an ENG LUBE LOW PFL after 15
seconds of straight, level, and unaccelerated
flight is a valid indication of low oil quantity.

If ENG LUBE LOW PFL was not present during fault
recall:7.

Land as soon as practical at the nearest
suitable airfield.

If ENG LUBE LOW PFL was present during fault
recall or recurs after fault recall:
8.

Stores-Jettison (if required).

9.

EPU switch-ON, if oil pressure decreases
below 10 psi.
Monitor hydrazine use. If consumption rateis too high, cycle EPU switch to OFF, thenNORM to conserve hydrazine. Be preparedto place EPU switch back to ON if the engine
seizes.

384Change 1

10. Land as soon as possible.
Plan to fly an SFO. Refer to SIMULATEDFLAMEOUT (SFO) LANDING and FLAMEOUT LANDING, this section.
11.

Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.

ENGINE FAULT Caution Light GE 129
Illumination of the ENGINE FAULT caution lightindicates that an engine PFL item was detected.
If ENGINE FAULT caution light illuminates:
1.

PFLD-Note PFL(s) displayed.

2.

C DF FACK, DR FAULT ACK button-Depress to acknowledge fault.

3.

Refer to PILOT FAULT LIST - ENGINE, thissection.

4.

C DF FACK, DR FAULT ACK button-Depress to perform fault recall.
The failure condition no longer exists if the
PFL is not present during the fault recall.

SEC Caution Light GE 129
Illumination of the SEC caution light indicates thatthe engine is operating in SEC. If the ENG CONT
switch is in  C DF PRI, DR NORM and the SEC
caution light illuminates, an automatic transfer to
SEC has likely occurred. Normal SEC operation ischaracterized by the SEC caution light, a fixed closed
exhaust nozzle, and AB unavailable. The ENG
CONT switch does not have to be positioned to SEC.
NOTE
If the rpm indication is also zero, theengine alternator has failed.
There are no throttle restrictions while operating
subsonic in SEC. If supersonic when transfer to SEC
occurs, the throttle must remain at MIL or above
until subsonic.
The thrust level at MIL, while operating in SEC, is7095 percent of that provided (for the same throttle
position and flight condition) in PRI. During landingin SEC, idle thrust is approximately twice that in PRI
with a normal nozzle because the nozzle is closed.

T.O. GR1F16CJ1
A throttle cable internal failure or a throttle linkage
disconnect results in an automatic transfer to SEC;
however, control of the engine with a failed throttle
cable or linkage disconnect is not possible in SEC. In
SEC, a throttle cable internal failure is character
ized by a possible rpm increase but no decrease withthrottle movements; a linkage disconnect results in
no response to throttle movements. In this situation,
control of the engine at idle and above can beregained in PRI by cycling the ENG CONT switch to
SEC, then back to  C DF PRI, DR NORM. The engine
logic for this failure is such that the engine remains
in PRI after cycling the ENG CONT switch. Afterlanding, shutdown of the engine must be accom
plished either with the FUEL MASTER switch or
maintenance personnel action to position the MEC
throttle input shaft to off.

When subsonic or if SEC caution light illuminates
when subsonic:
2.

Throttle-Verify engine responds normally tothrottle movement from IDLE to MIL; set as
desired.
AB operation is inhibited and exhaust nozzle
is closed.

If the engine is operating normally in SEC:3.

ENG CONT switch-Do not cycle.
The switch may remain in PRI or may beplaced to SEC. If the switch is placed to SEC,
do not place switch back to PRI.

If SEC caution light illuminates when supersonic:
1.

Throttle-Do not retard below MIL until
subsonic.

Retarding throttle below MIL while
supersonic may induce inlet buzz
which produces severe cockpit vibra
tion and probable engine stalls.

Cycling the ENG CONT switch in an
attempt to regain PRI may result inreoccurrence of the original malfunc
tion or a more severe condition.
4.

Land as soon as practical.
During landing in SEC, idle thrust isapproximately twice that in PRI with anormal nozzle.

Change 1384.1/(384.2 blank)

T.O. GR1F16CJ1
If the engine is operating abnormally in SEC:

When landing is assured (normally high key):

Failure to monitor sink rate and height
above terrain while applying low
thrust recovery procedures can resultin ejection outside ejection seat perfor
mance envelope.

Delaying engine shutdown can resultin a long, fast landing. Wheel braking
is less effective due to lack of WOW andthere is an increased probability of a
missed cable engagement.

3.

ENG CONT switch-Position to SEC, thenback to  C DF PRI, DR NORM.

4.

Airspeed-250 knots (If thrust is too low tosustain level flight).

5.

Land as soon as possible.
NOTE
A broken throttle cable or throttle
linkage disconnect causes a transfer to
SEC and abnormal engine response in
SEC. Reselecting PRI restores normal
engine operation for flight; however,engine shutdown after flight requires
either use of the FUEL MASTER
switch or maintenance personnel action
to position the MEC throttle input shaft
to off.

If thrust is too high to permit a safe landing:
NOTE
If throttle is stuck, control might beregained by depressing the cutoff
release, rotating the throttle outboard,and applying necessary force.
6.

Plan a flameout landing. Refer to FLAMEOUTLANDING, this section.

Do not start the JFS if engine seizure
has occurred or is anticipated or ifengine failure is a result of fuelstarvation. Starting the JFS may
result in no brake/JFS accumulator
pressure for the brakes.

7.

Throttle-OFF.
If throttle is stuck or engine does not respond,
shut down the engine with the FUELMASTER switch. At MIL, the engine flamesout in approximately 6 seconds. At IDLE, theengine flames out in approximately 45
seconds.

8.

HOOK switch-DN (if required).

The hook may miss the cable if theaircraft is not slow enough to compress
the MLG struts sufficiently to makeWOW or if forward stick pressure is
held.
FTIT Indicator Failure GE 129
The engine does not use FTIT as a control parameter.
However, routine missions should not be continued
since FTIT cannot be monitored.
Zero RPM/Erroneous RPM Indication GE 129
If the RPM indicator fails to zero or displays
erroneous RPM indication, it indicates either amalfunction within the indicator itself or within theengine alternator. Total alternator failure also
causes an automatic transfer to SEC. The ENGINE
warning light also illuminates for engine alternatoror RPM indicator failure. Engine transfer to SEC
due to certain partial or total alternator failures,
results in loss of most engine fault reporting. Partial
alternator failure may not result in automatictransfer to SEC. If the engine does not transfer to
SEC, fault reporting continues. Partial alternator
failure may also result in transfer to hybrid
operation with AB operation inhibited. In this case,
ENG HYB MODE and ENG A/B FAIL PFL's occur.
Routine missions should not be continued without afunctional RPM indicator.

Change 1385

T.O. GR1F16CJ1
If SEC caution light is illuminated:
1.

Go to SEC CAUTION LIGHT, this section.

If SEC caution light is not illuminated:
1.

Land as soon as practical.

Abnormal Engine Response GE 129
Refer to LOW THRUST ON TAKEOFF OR AT LOW
ALTITUDE (NONAB) GE 129 , this section, if
appropriate.
If thrust is too low to sustain level flight, turnimmediately toward the nearest suitable runway andestablish 250 knots airspeed. Consider jettisoning
stores to increase flying time available to complete
actions designed to restore usable thrust and improverange in the event those actions are unsuccessful.
NOTE
In a partial thrust situation, thrustavailable may increase as altitudedecreases. 250 knots approximates theairspeed at which thrust required forlevel flight is the lowest.
Abnormal engine response is varied and generallyindicated by abnormal thrust in relation to throttle
position, engine oscillations (either continuous,
momentary, or recurring), a complete lack of engine
response to throttle movement, autoacceleration/de
celeration, exhaust nozzle failure, or insufficientthrust. A PRI malfunction can cause these abnormal
engine responses as well as overtemperature oroverspeed indications. During AB operation, if engine
anomalies occur in region 1, the throttle should be
moved out of the AB range and AB should not be
selected again for the remainder of the flight.
Exhaust nozzle control or oil system malfunctions
can result in the nozzle being too far open, too far
closed, or unstable. Insufficient thrust may result if
the nozzle is more open than normal. SEC should beselected and, if the nozzle goes fully closed, engine
operation should be continued in SEC. If the nozzle
does not close in SEC, the ENG CONT switch should
be moved back to C DF PRI, DR NORM as morethrust results and AB lightoff may be attainable atlower altitudes. AB should only be used whenrequired to sustain flight.
If the exhaust nozzle is positioned more closed thannormal, thrust is adequate, but engine stall may
occur if AB is selected. Automatic transfer to SEC
should occur for most exhaust nozzle malfunctions.

386

NOTE
Certain PRI malfunctions may not
result in an autotransfer to SEC. If theengine operates abnormally, do not
rely upon an autotransfer to SEC.
Timely selection of SEC may preclude
further engine problems.
If thrust is still insufficient to make a safe landing
after selecting SEC or abnormal engine response isstill present, reattempt PRI.
If the throttle is stuck and thrust is suitable forsustained flight, attempts to free the throttle should
be delayed until within gliding distance of a suitablelanding field. If throttle is stuck in AB, placing the
ENG CONT switch to SEC terminates AB andprovides SEC MIL thrust. If throttle is stuck orotherwise prevented from normal movement,
control might be regained by depressing the cutoff
release, rotating the throttle outboard, and applying
necessary force.
If thrust is too high to permit a safe landing, use
excess thrust to climb and maneuver toward thenearest suitable airfield. Once high key for a
flameout landing is assured, follow procedures as
outlined in FLAMEOUT LANDING, this section.
Activate the JFS and EPU and then shut down theengine as soon as landing is assured (normally highkey) by placing the throttle to OFF or, if necessary,by placing the FUEL MASTER switch to OFF.
If abnormal engine response occurs:

FFailure to monitor sink rate and height

above terrain while applying low
thrust recovery procedures can resultin an ejection outside ejection seat
performance envelope.

FIf the throttle is stuck and thrust issuitable for sustained flight, attempts
to free the throttle should be delayed
until within gliding distance of asuitable landing field.

Idle PRI thrust with nozzle closed isapproximately 50 percent greater thanidle SEC thrust.

T.O. GR1F16CJ1
If in AB or if supersonic:
1.

5.

Throttle - Retard to MIL.

Do not start the JFS if engine seizure
has occurred or is anticipated or ifengine failure is a result of fuelstarvation. Starting the JFS may
result in no brake/JFS accumulator
pressure for the brakes.

Retarding throttle below MIL while
supersonic may induce inlet buzz
which produces severe cockpit vibra
tion and probable engine stalls.
If subsonic or problem still exists:2.

ENG CONT switch - SEC.

Plan a flameout landing. Refer to FLAMEOUTLANDING, this section.

When landing is assured (normally high key):

NOTE
Transfer to SEC may be accomplishedwhile supersonic if the throttle
remains at MIL.
3.

Airspeed-250 knots (if thrust is too low tosustain level flight).

4.

Throttle-Verify engine responds normally tothrottle movement from IDLE to MIL; set as
desired.
AB operation is inhibited and exhaust nozzle
is closed.

Delaying engine shutdown can resultin a long, fast landing. Wheel braking
is less effective due to lack of WOW andthere is an increased probability of a
missed cable engagement.
6.

Throttle-OFF.
If throttle is stuck or engine does not
respond, shut down the engine with the
FUEL MASTER switch. At MIL, the engine
flames out in approximately 6 seconds. At
IDLE, the engine flames out in approxi
mately 45 seconds.

7.

HOOK switch-DN (if required).

If a safe landing can be made with the current thrust:
5.

Land as soon as practical.
During landing in SEC, idle thrust is higher
than normal.

If thrust is insufficient to make a safe landing or
abnormal engine response is still present:
5.

ENG CONT switch- C

DF PRI, DR NORM.

NOTE
If the exhaust nozzle is failed open,
additional thrust to sustain flight may
be attained by selecting AB at low
altitudes.
6.

Land as soon as possible.

If thrust is too high to permit a safe landing:
NOTE
If throttle is stuck, control might beregained by depressing the cutoff
release, rotating the throttle outboard,and applying necessary force.

The hook may miss the cable if theaircraft is not slow enough to compress
the MLG struts sufficiently to makeWOW or if forward stick pressure isheld.
AB Blowout/Failure To Light GE 129
AB blowouts are characterized by a thrust loss, a fuelflow decrease, and nozzle closure. The DEC attempts
to relight the AB as long as the throttle remains inAB.
Refer to Figure 117. If AB is selected in region 1 andfails to light within 10 seconds or if the AB blows out
in any region, retard throttle out of AB. AB operation
should not be reattempted for the remainder of theflight.

Change 1387

T.O. GR1F16CJ1
AB operation in region 2 may include AB no lights ordelayed lights if unfavorable AB lightoff conditions
exist. If these are encountered, the throttle may beleft in AB. If the throttle is retarded, an AB light may
be attempted.
ENGINE STALLS GE 129
The three primary causes of a stall are inlet flow
distortion, AB instabilities, and hardware malfunc
tions. During normal aircraft operation, inlet flow
distortion severe enough to cause engine stall is notexpected. However, under some departure condi
tions, inlet flow distortion may induce engine stalls.
Hardwareassociated stalls may result from a failed
nozzle, control system malfunctions, or FOD.
The first indication of a stall at high thrust settings
may be a loud bang or pop. At lower thrust settings,
the first indication may be loss of thrust, lack ofthrottle response, or decreasing rpm. The throttle
should be retarded at the first indication of a stall.
Engine stalls are divided into two types:ABasso
ciated engine stalls and nonAB engine stalls.
ABAssociated Engine Stalls GE 129
Types of ABassociated engine stalls are:

S AB initiation - Stall at AB lightoff.
S AB sequencing - Stall during AB sequencing as theAB fuel flow increases with the throttle in AB.

S AB cancellation - Stall during throttle retard fromAB.

S AB blowout/relight - Stall occurs during relightafter a blowout in stabilized AB. May be preceded by
AB rumble.

ABassociated engine stalls are normally accompa
nied by a loud bang or pop and a possible fireball fromthe engine exhaust and occasionally the engine inlet.
These characteristics could be mistaken for anaircraft fire. Whenever a stall occurs while operatingin AB, the throttle should be snapped out of AB to
MIL. This usually clears the stall and restores
normal operation. The stalls may continue at MIL
and are characterized by bangs or pops of lower
intensity than AB stalls. If stall condition persists,refer to NONAB ENGINE STALLS GE 129 , thissection.
388

If an ABassociated engine stall occurs
below 30,000 feet MSL or while
supersonic, IGV system damage may
occur. Throttle movements after thestall clears should not be rapid.
FTIT fluctuation and decreasing rpm will probably
accompany stalls. The throttle should be retarded to
IDLE if the stall continues for a few seconds. The
engine may continue to stall at IDLE but may not be
audible, particularly at high altitudes. The engine
instruments should be monitored for indications ofstall. FTIT may rise while rpm decreases.
If the engine stalls at low altitude, an immediateclimb should be initiated. Retarding the throttle to
MIL may clear the stall. If engine response at low
altitude is not sufficient to maintain or gain altitude
and a suitable field is not immediately available,
ejection may be required.
NonAB Engine Stalls GE 129
NonAB engine stalls may occur if the control system
is malfunctioning, particularly during throttle
transients. NonAB engine stalls are often asymptom of a serious engine problem. NonABengine stalls may be inaudible; the first indication
may be a lack of engine response to throttle
movement which may be difficult to differentiate
from abnormal engine response. However, nonABengine stalls may be characterized by bangs, pops,
low intensity engine rumble or vibration, and/or
erratic orangeyellow flame from the engine
exhaust.
This exhaust flame could be mistaken for an engine
fire. FTIT fluctuation and decreasing rpm willprobably accompany stalls. If a stall is confirmed,
the throttle should be immediately retarded to
IDLE. The engine may continue to stall at IDLE but
the stall may not be audible, particularly at highaltitudes. The engine instruments should bemonitored for indications of stall. FTIT may
increase while rpm decreases. If the stall continues,
place ENG CONT switch to SEC. If the engine
recovers, throttle movements should be minimized
and made slowly until landing is assured.
If stall continues, initiate airstart. Refer to
AIRSTARTS GE 129 , this section. When trying toclear a stall with an airstart, the throttle should bemaintained in OFF for a few seconds to allow thestall to clear. If the stall continues after the airstart,
the engine may have a serious hardware problem.
The focus should shift to using available thrust toland at the nearest divert field.

T.O. GR1F16CJ1
If stall(s) occurred at MIL or below, or in AB below
30,000 feet MSL or while supersonic:
Prolonged engine operation with FTIT
in excess of 980_C can result insignificant engine damage and may
cause a nonrecoverable engine failure.
Engine Stall Recovery GE 129
If an AB stall(s) occurs:
1.

Throttle - Snap to MIL.

If AB stalls do not clear or stall(s) occurs below AB:
NOTE
NonAB stalls may be inaudible.
2.

Throttle-IDLE.

If stalls continue:
3.

ENG CONT switch-SEC.

If stalls continue:
4.

Throttle-OFF for a few seconds, then initiate
airstart. Refer to AIRSTART PROCEDURES
GE 129 , this section.
NOTE
For serious hardware problems, theengine may operate normally at idle
rpm but exhibit stall/vibration condi
tions at thrust settings above idle rpm.
Attempting additional airstarts will
not clear the condition. Use the highest
thrust setting below the stall/vibra
tion condition to sustain flight.

If stall(s) clears:
5.

Throttle-MIL or below. Minimize throttle
movements and make necessary movements
slowly.
NOTE
If stall(s) occurred in AB at 30,000 feet
MSL or above and while subsonic, theengine is safe to operate in the IDLE to
MIL range provided no other abnormal
engine indications are observed.

6.

Land as soon as possible.

INLET BUZZ GE 129
Inlet buzz occurs at supersonic airspeeds if an engine
control system failure or a CADC mach signal failure
results in insufficient airflow or if the throttle isretarded below MIL while operating in HYB or SEC.
Inlet buzz causes moderate to severe vibration
within the cockpit and may result in multiple engine
stalls.
If inlet buzz occurs, do not move the throttle until
subsonic. Decrease airspeed to subsonic as quickly
as possible by opening the speedbrakes andincreasing g. If engine stalls occur and persist, thethrottle should be retarded to IDLE when subsonic.
If the stalls do not clear, retard the throttle to OFF
for a few seconds, then advance to midrange. Refer
to AIRSTART PROCEDURES GE 129 , this section.
BIRD STRIKE GE 129
In the event of a bird strike or suspected bird strike,AB should be used only if absolutely necessary. It is
possible to lodge bird remains in the AB system suchthat liner damage and subsequent duct burn
through occurs if AB is used. There is no concern ofliner damage during any nonAB operation. Refer to
ABNORMAL ENGINE RESPONSE GE 129 , thissection, if appropriate.
ENGINE OVERSPEED GE 129
An overspeed occurs when engine rpm exceeds 109.5
percent. If an overspeed condition occurs, an MFL isrecorded and the engine control attempts to reduce
rpm below maximum limit. However, if the DEC
malfunctions and engine rpm reaches 113 percent,the overspeed protection in the MEC closes theoverspeed fuel shutoff valve resulting in a flameout.
To restore fuel, retard the throttle to OFF thenadvance to midrange. Refer to AIRSTART PROCEDURES GE 129 , this section.
ENGINE FAILURE OR FLAMEOUT GE 129
If the engine flames out, fuel starvation ormechanical failure has occurred.
A flameout is indicated by a decrease in FTIT andengine rpm decaying below inflight idle (approxi
mately 70 percent rpm). Loss of thrust and lack of
response to throttle movement confirm the flameout.
389

T.O. GR1F16CJ1The ENGINE warning light illuminates whenengine rpm is below 60 percent. Additionally, theMAIN GEN and STBY GEN lights illuminate below
50 percent rpm and the EPU should start running.
Do not mistake a loss of ECS noise as an engine
flameout.
A flameout indicates an engine control failure, fuelstarvation, fuel system malfunction, or fuel cutoff
due to engine overspeed protection. If the engine
flames out, two features may instantly restart theengine. There is an autorelight feature and thecapability to automatically transfer to SEC if certain
faults are detected in PRI. If these features work, therestart may take place instantly and the flameout
may not be noticeable (except for the illumination of
the SEC caution light). In this situation, remain in
SEC. (Refer to SEC CAUTION LIGHT, this section.)If the flameout progresses to the point that it isnoticeable, retard the throttle to OFF, then advanceto midrange. Refer to AIRSTART PROCEDURES
GE 129 , this section.

airstart or to additional glide range to reach asuitable landing field. Above 310 knots, more time isavailable by performing a zoom climb using a 3g
pullup to 30degree climb until approaching thedesired airspeed (use approximately 50 knots lead
point) and then initiating a zerog pushover. Below
310 knots, more time is available by performing a
constant altitude deceleration to the desired
airspeed; if required, climb to achieve minimum
recommended ejection altitude.
If the zoom results in an altitude below 4000 feet
AGL, there may be insufficient time to achieve anairstart prior to reaching minimum recommended
ejection altitude. In that case, primary consideration
should be given to preparing for ejection; do not delayejection below 2000 feet AGL.
If low altitude engine failure or flameout occurs:
1.

Zoom.

2.

Stores-Jettison (if required).
If stores jettison is attempted after main andstandby generators drop off line but before
EPU generator powers the SMS (approxi
mately 5 seconds delay), stores will notjettison.

Tower Shaft Failure GE 129
Failure of the engine tower shaft or its associatedgeartrain results in engine flameout due to fuelstarvation. A restart is not possible; primaryemphasis should be on a flameout landing. If unable
to make a flameout landing, refer to EJECTION(TIME PERMITTING), this section. Because tower
shaft failure results in the loss of rotation to theenginedriven gearbox and ADG, the initial symp
toms are similar to main fuel pump failure. The
primary differences are that the rpm indication drops
immediately to zero and ENGINE warning light andthe SEC caution light illuminate since the engine
alternator is no longer providing power to the DEC.
The JFS should be started immediately uponentering the JFS envelope to conserve EPU fuel. The
JFS drives the ADG and the engine gearbox which
restores rotation to both hydraulic pumps andprovides a reduced FLCS PMG output. Dependingon JFS performance and load, rpm may even be highenough to restore standby generator power;
however, main generator power may cycle on and off.
Without the load of the engine, the JFS produces a3055 percent rpm indication, which is the speed ofthe engine gearbox and not the actual engine rpm.
The true engine rpm is unknown.
Low Altitude Engine Failure or Flameout GE 129
Refer to figures 310 and 311. Initial reaction to any
malfunction at low altitude should be to trade excess
airspeed for altitude. Higher altitude translates
directly to either additional time to achieve an
390NOTE
Visually confirm the stores havejettisoned and jettison again if required.
3.

Perform airstart (if altitude permits). Refer to
AIRSTART PROCEDURES, this section.

Below 4000 feet AGL, there may beinsufficient time to perform an airstartprior to minimum recommended ejec
tion altitude.
AIRSTARTS GE 129
Refer to figure 312. Airstarting the engine does not
require exact airspeeds or rpm ranges, but there are
key events in the airstart sequence that must be
performed in a timely manner in order to have the best
chance for an airstart. The key events are initiating
the airstart while engine rpm is still high, selecting
SEC if there is no lightoff prior to rpm decaying below
50 percent in PRI (or immediately when below 10,000
feet AGL), and preserving engine rpm prior tolightoff.

Low Altitude Zoom Capability
DATA BASIS ESTIMATEDENGINE F110-GE-129

CONFIGURATION:

CONDITIONS:GW = 23,000-25,000 LB
DI = 0-50
LG UP

WINDMILLING OR SEIZED ENGINE
30-DEGREE CLIMB MAINTAINED TO 170/250 KIAS

12
3G PULLUP TO 30
ZOOM CLIMB TO
ACHIEVE:
10250 KIASINITIAL ALTITUDE

1000 FT AGL

170 KIAS

8

6

4

2

0

160

180

200

220

240

260

280

300

320

340

INITIAL AIRSPEED

360

KIAS

Figure 310.

380

400

420

440

460

480

500
1F-16X-1-4025X

T.O. GR1F16CJ1

391

0DATA BASIS ESTIMATEDENGINE F110-GE-129

CONFIGURATION:

CONDITIONS:GW = 23,000-25,000 LB
DI = 0-50
LG UP

30 DIVE TO DESCENT KIAS OR 3G PULLUP
TO 30 ZOOM CLIMB INITIATED FROM THE
AIRSPEED/ALTITUDE EXISTING AT FIRSTRECOGNITION OF ENGINE FAILURE

45 SECONDS ASSUMED AFTERTHROTTLE ADVANCE TO ACHIEVE
USABLE THRUST (ASSUMES AIRSTART
INITIATION AT 25 PERCENT RPM)

AIRSTART INITIATED AT START OF DIVE ORZOOM BY CYCLING THROTTLE TO OFF AND
THEN MIDRANGEDESCENT AIRSPEED IS 170 KIAS
(SEC) (JFS RUN LIGHT ON)

10
9

INITIAL ALTITUDE

1000 FT AGL

8
7
6
5
4
3
2
1
0

MINIMUM RECOMMENDED
EJECTION ALTITUDE

0

100

120

140

160

180

200

220

240

260

INITIAL AIRSPEED

280
KIAS

Figure 311.

300

320

340

360

380

400

420

440
GR1F-16CJ-1-0120A37

T.O. GR1F16CJ1

392

Low Altitude Airstart Capability

T.O. GR1F16CJ1Engine RPM and FTIT Response During
Spooldown and Airstart
ENGINE F110-GE-129

CONDITIONS:DRAG INDEX = 0

GW = 26,000 LB

KIAS = 250

NO WIND

POINT B TO
SL = 8.5 MIN/44 NM

501000 FEET40

ALTITUDE

250 KIAS(RECOMMENDED)

A
30
DIVE
B

170 KIAS30

400 KIAS/0.9 MACH
AIRSTART ENVELOPE
20JFS ENVELOPE

100

MIN KIAS(WITH JFSRUN LIGHT
ON)
0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

MACH NUMBER

OPTIMUM FLIGHT PATH DURING AN AIRSTART (TYPICAL)Engine out descent flight path maintains the aircraft in the required airstart envelope.
A 30-degree dive to descent KIAS is used to approach the airstart envelope(Point A to B).
Airstart is enhanced by increased airspeed, altitudes below 30,000 feet, and light-off
at highest possible engine rpm.

1F-16X-1-4024-1A

Figure312.(Sheet 1)

393

T.O. GR1F16CJ1

900

FTIT

C

800
700
600
500
LIGHT-OFF

400

LIGHT-OFF

300
200
100
90

RPM

%

80
70
60

LIGHT-OFF
LIGHT-OFF

50
40
30
20

0

10

20

30

40
TIME

50

60

70

80

90 0

10

SECONDS

20

30

40
TIME

RATE OF ENGINE RPM AND FTIT DECAY ANDRECOVERY FOR JFS-ASSISTED AIRSTART AT
20,000 FEET.
Following the rapid FTIT rise and peak after lightoff, FTIT slowly decreases approximately 50 C.
Airstarts initiated between 25-50 percent engine
rpm are slow to light-off and may take up to 90
seconds to regain usable thrust.

50

60

70

80

90

100

SECONDSRATE OF ENGINE RPM AND FTIT DECAY AND RECOVERY FOR SPOOLDOWN AIRSTARTS AT 25,000 FEET.
Following the rapid FTIT rise and peak after light-off, FTIT
slowly decreases approximately 50 C.
Airstarts initiated between 25-50 percent engine rpm areslow to light-off and may take up to 90 seconds to regain
usable thrust.

1F-16X-1-4024-2X

Figure312.(Sheet 2)

394

T.O. GR1F16CJ1
Oil pressure is directly related to rpm. Do notconfuse a low oil pressure indication due towindmilling rpm as an oil system malfunction.
Factors such as altitude, airspeed, weather, etc.,
must be considered in determining whether to try
an airstart, to accomplish a flameout landing, or toeject. Jettisoning stores reduces altitude loss
during an airstart and improves glide ratio during
flameout landing. If gliding distance is not a factor,
maintain 250 knots or more in order to reduce rpm
rate of decay until the JFS can be started. The
engine can be airstarted with airspeeds from170400 knots/0.9 mach; however, 250 knots
provides the best tradeoff of altitude loss, range,
and airflow for the engine.
In flight, the throttle must be retarded to OFF thenback to the operating range for only four reasons: toreset the overspeed protection logic, to clear a stall,
to begin the airstart procedure, or to terminate a
hot/hung start. Exact throttle position is notimportant for an airstart, so any position between
IDLE and MAX AB is acceptable; however, themidrange position is preferred because of possible
throttle misrigging at IDLE or possible engine
overspeed shutdown at MIL or above.
Once the throttle is retarded to OFF and then backto the normal operating range, do not retard thethrottle to OFF again during the airstart unless ahot/hung start occurs. Unnecessarily retarding thethrottle to OFF terminates any start attempt which
may be in progress.
A successful restart depends on many variables:
cause of flameout, type of fuel, altitude, airspeed,
and engine rpm when the airstart is attempted. Highengine rpm is the most important variable andprovides the best chance of a successful restart.
Therefore, do not delay the initiation of an airstartin an attempt to reach a particular flight condition.
Initiate the airstart as soon as it becomes apparent
that engine rpm has decayed below inflight idle
(approximately 70 percent rpm) or illumination of
the ENGINE warning light, engine instrument
indications, and no response to throttle movement
confirm a flameout. The best conditions for airstart
are below 30,000 feet MSL, at 250 knots or more, andwith high engine rpm.

At medium and high altitudes, the airstart attemptshould be started in the engine control mode selectedby the DEC. The DEC contains diagnostic logic
designed to identify PRI engine control failures andmay automatically transfer to SEC. If there is no
indication of a lightoff before rpm decays below 50
percent, place ENG CONT switch to SEC (even if the
SEC caution light is on) and continue the airstart
attempt. At low altitude (below 10,000 feet AGL),
SEC should be selected as soon as possible after
initiating the airstart.
Of equal importance to selecting SEC when required
is preserving engine rpm. The JFS should be started
as soon as the aircraft is in the JFS envelope. The
advantage of using the JFS to assist the airstart isthat once the JFS RUN light is on, airspeed can be
reduced. Under normal conditions the JFS will
motor the engine at a minimum of 25 percent.
An airstart can be rapid if lightoff occurs above 60
percent rpm. Airstarts initiated between 5025
percent engine rpm are slow to light off and may takeup to 90 seconds to regain usable thrust. If altitude
is available, increasing airspeed can assist engine
acceleration and decrease the time to regain usable
thrust once a lightoff is achieved. As long as engine
rpm continues to increase, this condition should not
be considered as a hung/no start. Spooldown
airstarts initiated below 25 percent rpm have beensuccessful during flight tests, but spool up to usable
thrust may take more time than is available. Keepengine rpm at 25 percent or above during spooldown
airstarts, if possible.
Following the rapid FTIT rise and peak of a lightoff,
FTIT slowly decreases approximately 50_C. There
fore, do not confuse a drop in FTIT as an unsuccessful
airstart unless accompanied by decreasing rpm as
well.
High Altitude Airstart Considerations GE 129
As altitude is increased above 30,000 feet MSL, theprobability of a successful airstart can be improved byattempting the airstart as soon as possible (before
rpm decays below approximately 50 percent) and by
quickly descending to altitudes below 30,000 feet
MSL after the airstart is initiated. Airspeeds above
250 knots (400 knots/0.9 mach maximum) should beconsidered as a means to reduce altitude and increase
the probability of a successful airstart. Spooldown
airstarts can be achieved with rpm as low as 25
percent, but not at all airspeeds and altitudes.

395

T.O. GR1F16CJ1
At high altitudes, dive as required to maintain speed
in the 250400 knot/0.9 mach range. Unless anairstart is obviously impossible (total lack of fuel,tower shaft failure, engine seizure, etc.), do not
become tempted to establish a maximum range or
maximum endurance glide. The first consideration
should be an immediate spooldown airstart attempteven if the engine fails for no apparent reason. If aspooldown airstart is not successful before reaching
20,000 feet MSL, a JFSassisted airstart should beattempted. When below 20,000 feet MSL, turn JFS
on. Activating the JFS above 20,000 feet MSL isprohibited since successful JFS start/motoring ofengine is unlikely and the brake/JFS accumulators
will be depleted. If the JFS RUN light is on, airspeed
may be reduced to achieve maximum range or
maximum endurance ( C 200 or 170, D 205 or 175
knots, respectively, plus 5 knots per 1000 pounds offuel/store weights). Time constraints due to EPU fuelconsumption must also be considered. A maximum
range or maximum endurance glide from above
approximately 35,000 feet MSL may exhaust EPU
fuel prior to landing. (Refer to T.O. GR1F16CJ11,
figure B63.) With the JFS running, EPU fuelconsumption is also reduced.
Low Altitude Airstart Considerations GE 129
Initiate the airstart as soon as possible. Afterinitiating a zoom climb and jettisoning stores (ifrequired), retard the throttle to OFF then advancethe throttle to the normal operating range. Place the
ENG CONT switch to SEC and turn on the JFS(START 2) to assist the airstart.
Following a zoom climb, plan to arrive at 250 knots
until the JFS RUN light is on; airspeed may then be
reduced to achieve maximum range or maximum
endurance ( C 200 or 170, D 205 or 175 knots,
respectively, plus 5 knots per 1000 pounds offuel/store weights). If a higher airspeed is maintainedor an attempt is made to gain airspeed to delay therpm decay, available time may be reduced to the pointthat an airstart is not possible.
During any low altitude airstart attempt, constantly
evaluate altitude above the ground relative to airstartsuccess. Do not delay ejection below 2000 feet AGL
unless the engine is producing thrust capable ofmaintaining level flight or safely controlling the sinkrate or unless a flameout landing can be accomplished.
Airstart Procedures GE 129
To begin the airstart sequence, retard the throttle toOFF; then immediately advance the throttle back into
the normal operating range, preferably midrange.

396Change 1

NOTE
If the throttle is retarded to OFF toclear a stall, it should be maintained in
OFF for a few seconds to allow the stall
to clear.
After throttle advance, monitor for signs of a lightoff
before rpm decays below 50 percent (characterized
by a rapid rise in FTIT accompanied by a slow
increase in rpm). If rpm and FTIT continue to decay
after rpm drops below 50 percent, place the ENG
CONT switch to SEC (even if the SEC caution lightis illuminated).
If a hot/hung start occurs, retard the throttle to OFF
and allow the FTIT to drop to below 700_C before
advancing the throttle. Increasing the airspeed(maximum of 400 knots/0.9 mach) should help thenext airstart to be cooler. If the condition persists,
retard the throttle to OFF, place the ENG CONT
switch to SEC, and allow the FTIT to decrease below
700_C before advancing the throttle.
After entering the JFS envelope, start the JFS toassist in preserving rpm. With the JFS RUN light on,
airspeed may be reduced to achieve maximum
range/endurance.
If the JFS stops running or fails to run within 30
seconds, do not reattempt a JFS start until thebrake/JFS accumulators have had time to recharge.
Allow 1 minute of engine rotation (either windmilling or JFSassisted) at 12 percent rpm or above toinsure that the brake/JFS accumulators are fullyrecharged. Recharging begins 34 seconds before the
JFS RUN light illuminates or 30 seconds after
selecting a start position (in the event of a JFS
failure to run). Recharging occurs regardless of JFS
switch position.
In the event of a JFS shutdown, the JFS switch does
not relatch in either start position while the JFS isspooling down. Spooldown from full governed speed
takes approximately 17 seconds. The JFS switch
must be cycled to OFF and then START 2 toreinitiate a JFS start. It is possible to complete thespooldown before the brake/JFS accumulators arerecharged if the JFS ran for only a short time.
It is possible the engine may not respond properly tothrottle movement following an otherwise success
ful airstart. If this occurs or if thrust is insufficient
to ensure a safe landing, switch to SEC, or if alreadyin SEC, switch to PRI.

T.O. GR1F16CJ1When the airstart is completed and usable thrust isregained, turn the JFS off. Reset the main generator
using the ELEC CAUTION RESET button andverify MAIN GEN and STBY GEN lights are off.
Cycle the EPU switch to OFF and then back to
NORM.

NOTE
FIf the JFS switch is erroneously placed

to START 1, leave it there.

FIf

the JFS RUN light does notilluminate or goes off once illuminated,
place the JFS switch to OFF andreattempt START 2 when the brake/
JFS accumulators are recharged. The
JFS switch does not relatch in either
start position while the JFS is spooling
down.

NOTE
The VHF radio is not powered when
the EPU is running.
To accomplish an airstart:
1.

Throttle-OFF, then midrange.

5.

Stores-Jettison (if required).
If stores jettison is attempted after main
generator drops off line but before EPU
generator powers the SMS (approximately 5
seconds delay), stores will not jettison.

FFTIT should decrease rapidly whenthrottle is OFF. If FTIT does notdecrease rapidly, verify that thethrottle is OFF.

NOTE
Visually confirm the stores havejettisoned and jettison again if required.

FDo not mistake a rapid initial FTIT riseduring an airstart as an indication of a
hot start. Typically, airstarts arecharacterized by rapidly increasing
FTIT with a slow increase in rpm.

If a relight does not occur before rpm decays below 50
percent or if below 10,000 feet AGL:
2.

ENG CONT switch-SEC (even if SEC
caution light is on).

3.

Airspeed-Attain approximately 250 knots or
establish maximum range or endurance
airspeed ( C 200 or 170, D 205 or 175 knots,
respectively, plus 5 knots per 1000 pounds offuel/store weights) with JFS RUN light on.
Above 30,000 feet MSL, airspeeds in the250400 knot/0.9 mach range should beconsidered to reduce altitude and increase
the probability of a successful airstart.
NOTE
If maximum gliding range is not a
factor, consider maintaining 250 knots
or more above 10,000 feet AGL toprovide best restart conditions (in caseof JFS failure). Below 10,000 feet AGL
with the JFS RUN light on, maintainmaximum range or maximum endur
ance airspeed.

4.

JFS switch-START 2 below 20,000 feet MSL
and below 400 knots.

If engine rpm rolls back or hangs below inflight idle
(approximately 70 percent) and FTIT exceeds 935_C:
6.

Throttle-OFF, then midrange.
Allow FTIT to drop below 700_C before
advancing the throttle.

7.

Airspeed-Increasemaximum).

(400

knots/0.9

mach

If hung start/hot start persists:
8.

Throttle-OFF.

9.

ENG CONT switch-SEC, if in PRI; PRI, if in
SEC.
NOTE
FIn certain circumstances, the engine

may malfunction in SEC. Returning to
PRI may result in attaining usable
thrust.

FThe proximity of the ENG CONT

switch to the JFS switch makes the
JFS switch susceptible to being
bumped to OFF when selecting SEC.

10. Throttle-Midrange.
Allow FTIT to drop below 700_C before
advancing the throttle.
Change 1397

T.O. GR1F16CJ1
If engine does not respond normally after airstart iscompleted:11.

ENG CONT switch-SEC.

12. Airspeed-250 knots (if thrust is too low tosustain level flight).
13. Throttle-Verify engine responds to throttle
movement; set as desired.
If engine does not respond normally after an airstartis complete in SEC, if thrust is still insufficient tomake a safe landing, or abnormal engine response isstill present:
14. ENG CONT switch- C

DF PRI, DR NORM.

15. Refer to FLAMEOUT LANDING, this section.
If engine responds normally:
16. JFS switch-OFF.
17. ELEC CAUTION RESET button-Depress.
Verify MAIN GEN and STBY GEN lights areoff.
18. EPU switch-OFF, then NORM.
19. ADI-Check for presence of OFF and/or AUX
warning flags.
If warning flag(s) is in view, refer to TOTALINS FAILURE, this section.

S Proximity of a suitable landing runway.
S Proficiency in performing simulated flameout(SFO) landings.

Due to the capabilities of the ejection seat, the entire
approach is within the ejection envelope; however,ejection should not be delayed in an attempt tosalvage a questionable approach. When performing
a flameout landing, the aircraft can safely stop (dryrunway without arresting gear) in approximately
twice the computed ground roll distance (8000footminimum runway length recommended), assuming
a touchdown no more than 1/3 of the way down the
runway at 1113 degrees AOA.
To perform a flameout landing, turn immediately
toward the desired runway. Jettison stores andestablish maximum range airspeed. Maximum
range airspeed may be less than the minimum
airstart airspeed. If range to the desired runway is
critical, the decision to attempt an airstart or aflameout landing rests with the pilot.
NOTE
FDuring an airstart attempt, do notslow below the minimum airstart
airspeed.

FIf the engine is still running, but thrustis insufficient to sustain level flight,
treat it as a flameout situation.

If only AUX flag is in view, pitch androll attitude information is likely to be
erroneous due to INS autorestart in
the attitude mode when other thanstraight and level, unaccelerated
flight conditions existed.
20. Land as soon as possible.
21. Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.
FLAMEOUT LANDING
The decision to eject or make a flameout landing
rests with the pilot. Considerations for attempting aflameout landing must include:

S Nature of the emergency.
S Weather conditions.
S Day or night.
398Change 1

Maximum range airspeed varies only with GW andis not affected by drag index. Maximum range
airspeed is C 200 knots for a GW of 20,000 pounds,
D 205 knots for a GW of 21,000 pounds, andincreases 5 knots per 1000 pounds of additional GW.
For most circumstances, sufficient accuracy is
obtained by adding 5 knots per 1000 pounds offuel/store weights PX III and by adding 5 knots if
CFT's are installed.
NOTE
FThis formula is based on the average

aircraft operating weight. Refer to T.O.
GR1F16CJ11, PART 1, DRAG IN
DEXES AND WEIGHTS-BASIC AIR
CRAFT. If range to desired runway is
critical, maximum range airspeed may
be calculated using actual GW in excess
of C 20,000 pounds, D 21,000 pounds.

FFor a 10,000foot descent (LG up), each

10 knots above or below maximum
range airspeed decreases glide range up
to 1/4 nm.

T.O. GR1F16CJ1
The maximum range airspeed equates to approxi
mately 7 degrees AOA (any GW or drag index) andprovides a glide ratio of approximately 7 nm per 5000
feet AGL (a no wind condition). Retaining stores or
flying into a headwind decreases glide range
significantly.
The EPU should be on and, if aircraft fuel isavailable, the JFS should be started using START 2
when below 20,000 feet MSL and below 400 knots
unless the engine is either seized or anticipated to
seize. The EPU should provide a minimum operationof 10 minutes (HYDRAZN light on) with normal
flight control demands before EPU fuel depletion.
Operating time can be extended to as much as 15
minutes if the JFS is running and flight control
inputs are minimized.
If expected time to landing exceeds expected EPU
operating time and excess energy is available, asteeper/faster descent may be flown. The JFS also
provides hydraulic pressure for normal braking and
NWS after landing.
When bleed air is no longer available to operate the
EPU, confirm that the EPU is operating on
hydrazine (EPU run and HYDRAZN lights on) since
the JFS alone does not provide adequate hydraulicpressure to land the aircraft. If the EPU isinoperative, maneuver the aircraft as necessary on
JFSassisted hydraulic pressure to a more favorable
ejection envelope and initiate ejection.
There are two basic types of flameout landing
patterns:the overhead approach (figure 310) or
the straightin approach (figure 311). The overhead
approach is preferred as it affords the mostopportunities to properly manage available energy
while providing the best visual cues for patterncorrections. The overhead approach may be enteredat any position, provided the proper altitude for thatpoint in the pattern can be obtained. The mainconcern is to reach high key, low key, or base key ator above the recommended minimum key altitudes.
A straightin approach is an alternate approach
when the overhead approach cannot be attained. For
both approaches, the initial aimpoint should beapproximately 1/3 of the way down the runway.
Overhead Approach
Refer to figure 313. Plan to arrive over the landing
runway (high key) at 700010,000 feet AGL. The
high key position may be approached from any
direction.

The recommended key altitudes are based on flying
a 360degree descending turn from high key with the
LG down. The altitudes vary with GW and withadditional drag due to stores. The recommended
high key altitude is C 7000, D 7500 feet AGL plus
500 feet per 1000 pounds of fuel/store weights PX III
and plus 500 feet if CFT's are installed. The
recommended low key altitude is C 3000, D 3250
feet AGL plus 250 feet per 1000 pounds of fuel/store
weights PX III and plus 250 feet if CFT's are installed.
These formulas include compensation for stores drageffects; thus, no additional correction is required.
If altitude will be significantly higher at high key,some form of altitude dissipating maneuver such as
a dive, gentle Sturns, or a 360degree descending
turn should be used. Speedbrakes also may be used
to lose excess altitude. However, if the speedbrakes
are not closed when a satisfactory flightpath isreached, the added drag may preclude a successful
flameout approach.
After departing high key, all attention should bedirected toward a successful landing. If actual
altitude at high key was below the recommended
altitude, fly maximum range airspeed with the LG
up until a satisfactory flightpath is reached and thenlower the LG. Optimum LG down airspeed is 10
knots less than maximum range (LG up) airspeed.
Minimum LG down airspeed is 20 knots less thanmaximum range (LG up) airspeed and provides
sufficient maneuverability to arrest the high sinkrate associated with a flameout approach. Optimum
angle of bank is 50 degrees with the LG up and 55
degrees with the LG down. Bank angles more than
10 degrees above/below optimum result in asignificant increase in altitude loss per degree ofturn and may preclude a successful flameout
approach.
NOTE
FDelaying LG extension until low key

allows successful completion of theoverhead approach from as low as 1500
feet below the recommended high key
altitude.

FAltitude loss for a 360degree descend

ing turn with the LG down increases
up to 500 feet for every 10 knots above
optimum LG down airspeed.

FAltitude loss for a 360degree descend

ing turn with the LG down increases
up to 500 feet for each 5 degrees
above/below the optimum bank angle.
Change 1399

T.O. GR1F16CJ1
The ground track of a flameout/SFO overhead
approach is approximately the same as that of anormal overhead approach except the final approach
is approximately 3/4 nm long. Avoid rapid flightcontrol inputs which use excessive EPU fuel and may
exceed the emergency hydraulic pump capability.

If EPU fuel quantity is below 25 percent
at high key (20 percent with the JFS
running), a flameout landing should
not be attempted since adequate
hydraulic pressure may not be avail
able through the landing.
StraightIn Approach
Refer to figure 314. If one of the overhead approachkey positions cannot be reached, a straightinapproach may be flown. The clean glide at maximum
range airspeed should be continued until the initial
aimpoint is 1117 degrees below the horizon; then the
LG should be lowered. Seventeen degrees is below theforward field of view. A good visual reference for 15
degrees is when the initial aimpoint is at the bottomof the HUD (just above the radome). Optimum LG
down airspeed is 10 knots less than maximum range(LG up) airspeed. Minimum LG down airspeed is 20
knots less than maximum range (LG up) airspeed andprovides sufficient maneuverability to arrest the highsink rate associated with a flameout approach.
NOTE
For a 10,000foot descent (LG down),
each 10 knots above optimum LG down
airspeed decreases glide range up to
1/2 nm.
IMC Penetration
Should IMC be encountered during a flameout
approach to the intended runway and no alternate
runway is available, an alternate descent/penetration may be flown which should allow maneuvering
airspeed after penetrating the undercast.

3100

IMC penetration should not be
attempted unless present position is
known and navigation can be per
formed throughout the descent, andhigh terrain or other hazards are not a
factor.
The stores should be jettisoned and the aircraft glided
at maximum range airspeed until a 1:1 ratio between
altitude in thousands of feet and range to the runway(e.g., 20,000 feet AGL at 20 nm, 15,000 feet AGL at 15
nm, etc.) is attained. The descent angle should thenbe increased and airspeed allowed to increase tomaintain the 1:1 ratio. This equates to a 910 degree
descent angle. This 1:1 glide ratio must be
maintained until sufficient airspeed is attained tomaneuver after penetrating the undercast.
NOTE
A 90 degree level turn at 50 degrees
bank angle with the LG andspeedbrakes retracted will dissipate
6585 knots. A 180 degree turn will
dissipate 145250 knots. Airspeed
dissipation increases with increasing
GW and DI. A glide angle at a 1:1 ratio
begun from maximum range airspeed
will result in an airspeed of 260320knots after a 10,000foot descent.
Higher airspeed at the start of theglide, additional descent altitude,
heavier gross weight, or lower dragindex will result in higher airspeed at
the completion of the glide.
At 3000 feet AGL, the aircraft should be 3 nm from thetouchdown point. If the runway is not in sight by base
key altitude, the aircraft may be zoomed for acontrolled ejection. When VMC is attained and the
runway is in sight, the aircraft should be glided to an
attainable key position for an overhead approach or
to a straightin approach and the LG should belowered. Excess airspeed above optimum LG down
airspeed not required to maneuver to the flameout
landing approach should be dissipated by use of
speedbrakes or early LG extension.

T.O. GR1F16CJ1Flameout Landing Pattern (Typical)(OVERHEAD APPROACH)NOTES:
1. Jettison stores (if required).

4. Altitudes:
High Key
7000-10,000 feet AGL
Recommended altitude is C 7000, D
7500 feet AGL plus 500 feet per 1000
pounds of fuel/store weights PX III and
500 feet if CFT's are installed.
50

2. Maximum range (LG up) airspeed is C 200, D
205 knots. Optimum airspeed (LG down) is C
190, D 195 knots. Minimum LG down airspeed
is C 180, D 185 knots. Increase airspeeds by
5 knots per 1000 pounds of fuel/store weights
PX III and an additional 5 knots if CFT's are
installed.
3. Maximum range (LG up) airspeed equates to approximately 7 degrees AOA (any GW or drag index) andprovides a glide ratio of approximately 7 nm per
5000 feet AGL. If stores are retained, glide ratio
decreases.

Low Key
3000-5000 feet AGL
Recommended altitude is C 3000, D
3250 feet AGL plus 250 feet per 1000
pounds of fuel/store weights PX III and
250 feet if CFT's are installed.
Base Key
2000 feet AGL minimum
5. Optimum bank angles are 50 degrees (LG up) and 55
degrees (LG down) for least altitude lost per degreeof turn.
A

B
C

A.

HIGH KEY

Above a point approximately 1/3 of the way down the runway

B.

LOW KEY

Abeam point of rollout on final

C.

BASE KEY

Midpoint of turn from downwind to final
Point of intended
touchdown

3/4 nmA
X

X

Initial aimpoint shouldbe 1/3 runway lengthC

1 nm
Low key position
varies with winddirection/velocity

B
GR1F-16CJ-1-0123X37

Figure 313.(Sheet 1)

Change 13101

T.O. GR1F16CJ1Flameout Landing Pattern (Typical)(OVERHEAD APPROACH)NOTES:
6. PX II With FCC off, HUD continues to computeflightpath marker and to position scales for use
during flameout approach.
7. Frost or condensation on the canopy could restrict
visibility during flameout approach. Place AIR
SOURCE knob to RAM and place DEFOG leverforward below 25,000 feet MSL.
8. Time constraints due to EPU fuel consumption must
be considered as well as distance to be covered.
To estimate required EPU fuel for a nonstandard
approach, use 15 percent per minute as a basis
for computation.

The JFS alone does not provide adequate
hydraulic pressure to land the aircraft.
Do not allow airspeed to decrease below
minimum LG down airspeed.
Eject if it becomes obvious that a safe
landing cannot be made. Ejection can beaccomplished at any point in the pattern; however, do not delay ejection
below 2000 feet AGL in an attempt
to salvage a questionable approach.

9. Starting JFS reduces load on EPU, conserves EPU
fuel, and partially restores hydraulic system B.
10. If alternate LG extension is used, the NLG may notindicate down until airspeed is reduced below
190 knots.

HIGH KEY
Do not extend LG unlessbase key is assured.

EPU fuel quantity should be at least 25
percent (20 percent with JFS running)at high key to insure adequate hydraulicpressure throughout landing.
ROLLOUTDrag Chute
DEPLOY(if required).
Speedbrakes
Open.
FLARE
Hook
DN (if required).
Touch down 11-13 degrees
AOA optimum. Speedbrakes
as required.

BASE KEY
LG down. Increase airspeed
and/or open the speedbrakes
to move touchdown closer toapproach end of runway.

Do not delay lowering LG
below 2000 feet AGL.

LOW KEY
Do not extend LG unlessbase key is assured.
GR1F-16CJ-1-0124A37

Figure 313.(Sheet 2)3102Change 1

T.O. GR1F16CJ1Flameout Landing Pattern (Typical)(STRAIGHT-IN APPROACH)NOTES:
6.

1. Jettison stores (if required).
2. Maximum range (LG up) airspeed is C 200, D
205 knots. Optimum airspeed (LG down) is C
190, D 195 knots. Minimum LG down airspeed
is C 180, D 185 knots. Increase airspeeds by
5 knots per 1000 pounds of fuel/store weights
PX III and an additional 5 knots if CFT's are
installed.
3. Maximum range (LG up) airspeed equates to approximately 7 degrees AOA (any GW or drag index) andprovides a glide ratio of approximately 7 nm foreach 5000 feet AGL. If stores are retained, glide ratio
decreases.
4. Minimum altitudes are based on an LG up glide at
maximum range airspeed to 2000 feet AGL followedby an LG down glide at optimum LG down airspeed
to the runway for a drag index of 100.

PX II With FCC off, HUD continues to computeflightpath marker and to position scales for use
during flameout approach.

7. Frost or condensation on the canopy could restrict
visibility during flameout approach. Place AIR
SOURCE knob to RAM and place DEFOG leverforward below 25,000 feet MSL.
8. Time constraints due to EPU fuel consumption must
be considered as well as distance to be covered.
To estimate required EPU fuel for a nonstandard
approach, use 15 percent per minute as a basis
for computation.
9. Starting JFS reduces load on EPU, conserves EPU
fuel, and partially restores hydraulic system B.
10. If alternate LG extension is used, the NLG may notindicate down until airspeed is reduced below
190 knots.

5. After lowering LG, glide range decreases by approximately 30 percent. Airspeed greater than optimum
LG down airspeed significantly increases energy
loss rate and decreases glide range.

ROLLOUTDrag Chute
DEPLOY(if required).
Speedbrakes
Open.
Hook
DN (if required).

FLARE
Touch down 11-13 degrees
AOA optimum. Speedbrakes
as required.
AREA C
4-0 nm(no wind)
4 nm (no wind)
4000-8000
feet AGL
8 nm (no wind)7000 feet AGL

POINT B

POINT A
GR1F-16CJ-1-0125A37

Figure 314.(Sheet 1)

Change 13103

T.O. GR1F16CJ1Flameout Landing Pattern (Typical)(STRAIGHT-IN APPROACH)

Do not allow airspeed to decrease below
minimum LG down airspeed.

EPU fuel quantity (points A, B1, and B 2 )should be sufficient to insure adequate
hydraulic pressure through landing.

If the aimpoint on the runway moves up in thefield of view while maintaining maximum range(LG up) airspeed, the runway probably cannot
be reached. This path corresponds to a glide
angle of about 7 degrees between the horizon
and the aimpoint.

Eject if it becomes obvious that a safe landing
cannot be made. Ejection can be accomplishedat any point in the approach; however, do notdelay ejection below 2000 feet AGL in an
attempt to salvage a questionable approach.

The JFS alone does not provide adequate
hydraulic pressure to land the aircraft.

*17*8

POINT A

*11
POINT B1

POINT B2

OVERHEAD
APPROACH

AREA CPOINT B
7000
FEET
AGL

2000
FEET
AGL

* Angle between horizon and aimpoint

POINT A

POINT B

AREA C

8 nm (no wind), 7000 feet AGL,
continue glide until initial aimpoint is 11-17 degrees below
horizon. Then lower LG andestablish optimum LG down
airspeed. As a guide, no windminimum EPU fuel is 45 percent(40 percent with JFS running).

4 nm (no wind), 4000-8000 feet
AGL, airspeed and LG as required.
As a guide, no wind minimum EPU
fuel is:4 - 0 nm (no wind), initial aimpoint
is more than 17 degrees below horizon (under nose of aircraft and not
visible). Normal straight-in approachis not feasible.

POINT B 1
6 nm
35 percent (30percent with JFS running).

Options are:

POINT B 2

Delay LG lowering. Plan an overhead approach from a high key
altitude but below the normal
recommended altitude.

4 nm
25 percent (20percent with JFS running).

Delay LG lowering. Plan a modified flightpath to low key.
Lower LG, open speedbrakes,
and dive and maneuver aircraft
to intercept a point on thenormal straight-in glidepath.

Do not delay lowering LG
below 2000 feet AGL.
GR1F-16CJ-1-0126X37

Figure 314.(Sheet 2)3104

T.O. GR1F16CJ1Landing Phase
The LG should be lowered no later than 2000 feet
AGL to allow adequate time for alternate LG
extension. Establish a glidepath to achieve the initial
aimpoint while maintaining optimum LG down
airspeed. Once wings level on final approach, beaware of the tendency to slow below minimum LG
down airspeed.

have chocks installed or engage parking brake.
Leave the battery on line until chocks are installed.
If JFS START 2 was attempted but was unsuccess
ful, no braking is available for stopping or directional
control unless the brake/JFS accumulators arerecharged. Use flaperons and rudder as required tomaintain directional control. As the aircraft slows
below 70 knots, directional control is reduced andthe aircraft may drift right.
Flameout Landing Procedures

Do not attempt to stretch a glide byallowing the airspeed to decrease
below minimum LG down airspeed. A
slower airspeed decreases the maneu
verability available to arrest the highsink rate associated with the flameout
approach and may preclude a success
ful flameout landing.
Once landing is assured, the recommended procedure
is to shift the aimpoint from 1/3 of the way down the
runway to a position short of the intended touchdown
point. Speedbrakes may be used to help control
airspeed. The higher the airspeed, the shorter the
aimpoint should be to allow for additional float (fromflare to touchdown). The aircraft is easiest to control
in the flare if the flare is begun between optimum andminimum LG down airspeeds. The point at which theflare is begun depends upon airspeed, sink rate, andglide angle. The flare should be started high enoughto allow a smooth gradual reduction in glide angle but
not so high as to run out of airspeed prior totouchdown. Under a no wind condition, the aircraft
floats 30004000 feet after beginning the flare, if theflare is begun at the optimum LG down airspeed.
Once the sink rate is arrested, attempt to slow to anormal touchdown airspeed and AOA. If excess
airspeed exists after arresting the sink rate, the best
method to slow the aircraft is to stay airborne until
normal touchdown airspeed is reached.
After Touchdown
After touchdown from a flameout landing, use anormal or short field stopping technique as requiredby the stopping distance available. If the JFS andEPU are running, normal braking and NWS areavailable (NWS is inoperative if the LG was loweredwith the alternate LG system). If the JFS is notrunning, only the brake/JFS accumulators areavailable to supply hydraulic pressure for braking.
Stop the aircraft by making one steady brakeapplication just short of antiskid cycling. If there isany doubt about stopping on the remaining runway,
lower the hook. When the aircraft is fully stopped,

If the engine has flamed out or if flameout isimminent, turn toward a suitable runway andaccomplish either an overhead approach or astraightin approach, as appropriate.

S Altitudes (overhead approach):
S High key-700010,000 feet AGL.
Recommended altitude is C 7000, D

7500 feet
AGL plus 500 feet per 1000 pounds of fuel/store
weights PX III and plus 500 feet if CFT's are
installed.

S Low key-30005000 feet AGL.
Recommended altitude is C 3000, D

3250 feet
AGL plus 250 feet per 1000 pounds of fuel/store
weights PX III and plus 250 feet if CFT's are
installed.

S Base key-2000 feet AGL minimum.

S Altitudes (straightin approach):S 8 nm-7000 feet AGL minimum.

The minimum altitude is based on an LG up
glide at maximum range airspeed to 2000 feet
AGL followed by an LG down glide at
optimum LG down airspeed to the runway fora drag index of 100. A lower drag index
slightly reduces the minimum altitude
required. A higher drag index slightly
increases the minimum altitude required.

S 4 nm-40008000 feet AGL.

Delay lowering the LG until the initial
aimpoint is 1117 degrees below the horizon.

Eject if a safe landing cannot be made.
Ejection can be accomplished at any
point in the pattern but do not delayejection below 2000 feet AGL in an
attempt to salvage a questionableapproach.
Change 13105

T.O. GR1F16CJ1
1.

Stores-Jettison (if required).

2.

Airspeed- C 200, D 205 knots.
Increase airspeed 5 knots per 1000 pounds offuel/store weights PX III and plus 5 knots if
CFT's are installed. This airspeed equates toapproximately 7 degrees AOA.

3.

JFS accumulators are recharged. The
JFS switch does not relatch in either
start position while the JFS is spooling
down.
5.

AIR SOURCE knob-RAM (below 25,000 feet
MSL).

NOTE

6.

DEFOG lever-Forward.

During an airstart attempt, do not slow
below the minimum airstart airspeed.

7.

LG handle-DN. (Use DN LOCK REL button
if required.)

EPU switch-ON.
NOTE
FDo not delay lowering LG below 2000

PX II The VHF radio is not powered byfeet AGL.

the EPU.
4.

FIf LG handle does not lower, selectJFS switch-START 2 below 20,000 feet MSL
and below 400 knots.

FEPU fuel quantity should be at least 25

BRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND. PW 229
Nozzle remains closed, resulting inhigher than normal landing thrust.

8.

percent (20 percent with JFS running)at high key for an overhead approachor 45 percent (40 percent with JFS
running) at 8 nm for a straightinapproach to insure adequate hydraulicpressure through landing.
JFS alone does not provide
adequate hydraulic pressure to land
the aircraft.

ALT GEAR handle-Pull (if required) (190knots maximum, if practical).
Alternate LG extension can be used up to 300
knots; however, the NLG may not fully extend
until 190 knots. Time above 190 knots should
be minimized in case there is a leak in thepneumatic lines.

FThe

FNWS is not available following alter

nate LG extension.

FDo not start the JFS if engine seizure

has occurred or is anticipated or ifengine failure is a result of fuelstarvation. Starting the JFS may resultin no brake/JFS accumulator pressure
for the brakes.
NOTE

FIf engine is not operating, consider

placing the FUEL MASTER switch to
OFF if a fuel leak exists. This action
may conserve fuel for the JFS.

FDo not depress the ALT GEAR reset

button while pulling the ALT GEAR
handle. This action may preclude
successful LG extension.

9.

Airspeed- C 190, D 195 knots optimum inpattern.
Increase airspeed 5 knots per 1000 pounds offuel/store weights PX III and plus 5 knots if
CFT's are installed.

FIf the JFS is erroneously placed to

START 1, leave it there.

FIf

the JFS RUN light does notilluminate or goes off once illuminated,
place the JFS switch to OFF andreattempt START 2 when the brake/

3106Change 1

Do not allow airspeed to decrease
below C 180, D 185 knots plus 5 knots
per 1000 pounds of fuel/store weights
PX III and plus 5 knots if CFT's are
installed.

T.O. GR1F16CJ1
After touchdown:

9.

10. DRAG CHUTE switch-DEPLOY (if required).
11.

SJ PAGE (MFD)-Select stores desired forjettison.

HOOK switch-DN (if required).
FJettisonof an inboard shoulder
mounted store from a TER at station 4
or 6 with the MLG down may result in
LG and store(s) collision. To avoid this,select RACK for jettison instead of
WPN.

If brake/JFS accumulator braking is used:
12. Stop straight ahead and engage parking
brake.

FJettison of external wing fuel tankswith stores/suspension equipment at
stations 3 and/or 7 with MLG down
may result in LG and external wing
fuel tank collision.

FBrakes should be applied in a single,

moderate, and steady application
without cycling the antiskid.

FFailure to load theactual storesconfiguration into SMS inventory could
cause damage to the aircraft byinhibiting the selective jettison releasetime delay used to insure safe
370/600gallon fuel tank separation
when a store is present at station 3 or 7.

FBrake pedal deflection of 1/16 inch

activates the brakes and bleeds thebrake/JFS accumulators. To avoid
brake activation and loss of accumula
tor fluid, do not rest feet on the brake
pedals.

FDo not attempt to taxi clear of theFSelective jettison airspeed/mach lim

runway. Loss of brake/JFS accumula
tor pressure results in the inability tostop or steer the aircraft.

its in T.O. GR1F16CJ12 are only
valid for:D Selective jettison of one store type at

13. Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.

a time.

D Selective jettison from nonadjacentstations.

JETTISON

If simultaneous selective jettison ofeither more than one store type or fromadjacent stations is required, adhere
to emergency jettison airspeed/machlimits.

Selective Jettison
Selective jettison is used to release selected store(s)(except airtoair missiles) or suspension equipment
in an unarmed or unguided condition.
1.

GND JETT ENABLE switch-ENABLE (if LG
is down).

2.

MASTER ARM switch-MASTER ARM.

3.

DR ARMT CONSENT switch-On.

4.

PX II SMS switch-SMS.

5.

PX III ST STA switch-ST STA.

6.

DOG FIGHT switch-Center.

7.

MFD-SMS format.
D Store and station selections can be madefrom either cockpit.

8.

SJ OSB (MFD)-Depress.

NOTE
FWeapon(s) and/or rack(s) to be jetti

soned is highlighted.

FWhen 300gallon and 370/600gallon

fuel tanks are carried simultaneously,
the 300gallon fuel tank must be
separated prior to the 370/600gallon
fuel tanks.

10.

WPN REL or
ton-Depress.

C

PX II

ALT REL butNOTE
When jettisoning tanks from stations 4
and 6, hold release button depressedfor 1 second.
Change 13107

T.O. GR1F16CJ1Emergency Jettison
Emergency jettison is a onestep operation whichclears all expendable stores and racks except
airtoair missiles. All weapons are released in an
unarmed or unguided condition. PX II If the SMS isoff, depressing the EMER STORES JETTISON
button supplies electrical power to the SMS. In
addition, while the button is depressed, the SMS
enters the EJ mode. When the button is released,
the avionic system returns to the previous operating
mode. Emergency jettison is not available unless themain, standby, or EPU generator is operating. PX III
Emergency jettison may not be available if an MMC
FAIL PFL message is present. In this case,emergency jettison can be restored by placing the
MMC switch to OFF.
PX III If the MMC is off, depressing EMER STORESJETTISON button supplies electrical power to the
MMC and RIU's. In addition, while the button isdepressed, the MMC enters the EJ mode.

resets. Caution level failures (except FLCS BUS
FAIL) also trigger the FLCS FAULT caution light;
this light goes off if the failure resets or when thefault is acknowledged. Refer to PILOT FAULT
LIST-FLCS, this section, for PFL list.
The FLCS emergency procedures are based on thefollowing recommended sequence of pilot actions:

S Note PFL(s) displayed.
S Acknowledge fault.
S Refer to PILOT FAULT LIST-FLCS, this section.
S Refer to appropriate FLCS emergency procedure.
S Perform FLCS reset, if so directed in the emergency
procedure.

S Perform fault recall.

1.

GND JETT ENABLE switch-ENABLE (ifrequired).
Use EMER STORES JETTISON on the
ground only as a last resort.

Certain MUX bus communication failures may resultin the inability to display FLCS PFL's on the PFLD.
If this occurs, PFL's may be observed on the FLCS
page of the MFD.

2.

EMER STORES JETTISON button-Depress(1 second).

Manual selection of DBU to attempt correction of aperceived FLCS failure is not recommended.

PX III Emergency jettison is notavailable if an MMC FAIL PFL
message is present. Emergency jetti
son can be restored by placing the
MMC switch to OFF.

NOTE
If the initial actuation of the EMERSTORES JETTISON button fails tojettison all aircraft stores, subsequentattempts may successfully release theremaining stores.
FLCS FAILURES
Significant FLCS failures, except for reversion to
DBU, cause one or more PFL's to be displayed.
Warning level failures also trigger the FLCS
warning light; this light goes off only if the failure

3108

Do not select DBU if the FLCS warning
light is on since doing so allows failed
signals which were previously voted
out to be used.
Air Data Malfunctions
A single failure in the air data system (static/impact
pressure) is indicated by the FLCS FAULT caution
light and an FLCS ADC FAIL PFL. No degradationof aircraft performance should be noticed.
Contamination of the pitot probe can cause blockageof the passage which provides total pressure signals.
This causes erroneously low airspeed indications in
the HUD and on the airspeed/mach indicator with an
FLCS ADC FAIL PFL at approximately 300 knots.
An erroneously low airspeed value is used for FLCS
gain scheduling which may result in pitch
oscillations.

T.O. GR1F16CJ1
If erroneously low airspeed indications are present
both in the HUD and on the airspeed/mach indicatorduring takeoff, consider aborting. If takeoff iscontinued, maintain airspeed below 210 knots (CatI) or 275 knots (Cat III) (use INS groundspeed or
wingman indications to determine airspeed) andland as soon as practical. Pitch oscillations may be
experienced at higher speeds. If oscillations areencountered, minimize stick commands and slow toless than 210 knots (Cat I) 275 knots (Cat III).
A dual failure of static or impact pressure systems isindicated by the FLCS warning light and a STBYGAIN PFL. The FLCS ADC FAIL PFL is still present
since the first failure is still active. In standby gains,the FLCS has transitioned to a fixed set of control
gains which are optimized for sea level and 600 knots(LG handle in UP) and sea level and 230 knots (LGhandle in DN). The LEF's are at zero degrees with the
LG handle up and ALT FLAPS switch in NORM, andare at 15 degrees down with the LG handle down or
the ALT FLAPS switch in EXTEND.
An FLCS reset can always be attempted for a first air
data failure. If the first air data failure does not reset,
and a second failure occurs (engaging standby gains),
the first failure is latched and will not reset. However,the second failure is allowed to be reset and standby
gains can be exited if the reset is successful.
Flying qualities in standby gains are degraded.
Airspeed should be maintained subsonic and lessthan 650 knots. AOA should be maintained less than
12 degrees. Since the gains are optimized for 600
knots at sea level, airspeed should be maintained
greater than 240 knots until the LG is lowered.
If FLCS ADC FAIL PFL occurs:
If BRK PWR DEGR PFL is also present, refer to
FLCS SINGLE ELECTRONIC FAILURE, thissection.
1.

Establish 1g flight.

2.

FLCS RESET switch-RESET.

If STBY GAIN PFL occurs:
1.

Establish 1g flight with maximum of 12
degrees AOA.
Airspeed 240650 knots with LG up.

2.

FLCS RESET switch-RESET.

3.

Land as soon as practical.
Do not slow below 240 knots with LG up if
STBY GAIN PFL is still present.

AOA Malfunction
A single failure in the AOA system is indicated byillumination of the FLCS FAULT caution light and an
FLCS AOA FAIL PFL. No degradation in aircraft
handling characteristics should be encountered. The
first AOA failure signal defaults to 11 degrees toprevent hardover commands if two AOA inputs fail atthe upper AOA limit.
A dual AOA failure illuminates the FLCS warning
light and triggers an FLCS AOA WARN PFL. The
FLCS system code and FLCS AOA FAIL PFL are stillpresent since the first failure is still active. With adual AOA failure, the FLCC selects the mid AOA
value among the remaining good signal, 11 degrees,
and the failed signal. Aircraft handling characteris
tics may be degraded somewhat in the cruise
configuration, but should be adequate for landing
since the first AOA failure defaulted to 11 degrees.
FLCS RESET may be attempted for any AOA failure.
If the first failure does not reset and a second failure
occurs, the first failure is latched and cannot be reset.
If FLCS AOA FAIL PFL occurs:
If BRK PWR DEGR PFL is also present, refer to
FLCS SINGLE ELECTRONIC FAILURE, thissection.
1.

Establish 1g flight.

2.

FLCS RESET switch-RESET.

If failure indications go off:3.

Continue normal operation.

If failure indications remain on:
If failure indications go off:3.

Continue normal operation.

If failure indications remain on:
3.

3.

Land as soon as practical.

Land as soon as practical.
Do not exceed 11 degrees AOA during
approach, landing, or twopoint aerodynamic
braking.

If FLCS AOA WARN PFL occurs:
1.

Establish 1g flight.
3109

T.O. GR1F16CJ1
2.

FLCS RESET switch-RESET.

3.

If FLCS warning light goes off:3.

S C PW 229 135, GE 129 136 knots

Land as soon as practical.
Do not exceed 11 degrees AOA during
approach, landing, or twopoint aerodynamic
braking.

S D PW 229 137, GE 129 138 knots
S Add 4 knots/1000 pounds of fuel/store
weights equates to 13 degrees AOA (add 8
knots for 11 degrees AOA).

If FLCS warning light remains on:
3.

Land as soon as possible.
Do not exceed 11 degrees AOA during
approach, landing, or twopoint aerodynamic
braking.

CADC Malfunction
A failure of any of the CADC electrical inputs (AOA,
total temperature, altimeter barometric reference,etc.) or a detected failure internal to the CADC causes
illumination of the CADC caution light. If the CADC
caution light does not reset, systems dependent on
CADC information should be checked for proper
operation.

Retarding the throttle below MIL
while supersonic may induce inlet
buzz which produces severe cockpit
vibration and probable engine stalls.
NOTE
If a CADC malfunction occurs, the FLCC
AOS feedback function may deactivate.
If CADC caution light illuminates:
1.

FLCS RESET switch-RESET.

Land as soon as practical.
Final approach airspeeds:

Servo Malfunction
An initial servo malfunction results in illumination of
the FLCS FAULT caution light and an ISA
LHT/RHT/LF/RF/RUD FAIL PFL for the applicable
ISA. A failure of the second servo in the same ISA
results in a repeat of the PFL. Multiple servo
malfunctions involving more than one ISA cause an
ISA ALL FAIL PFL. Servo failures are reset byplacing the FLCS RESET switch to RESET. No
degradation in flying qualities is expected if the servo
does not reset.
Hydraulic failures or momentary drops in hydraulicpressure (e.g., wake turbulence encounter, air inhydraulic system) also illuminate the FLCS FAULT
caution light and cause an ISA ALL FAIL PFL.
1.

Airspeed-400 knots maximum (subsonic).

If hydraulic failure is confirmed:
2.

Go to SINGLE/DUAL HYDRAULIC FAILURE,
this section.

If hydraulic pressures are normal:
3.

FLCS RESET switch-RESET.

If failure indications go off:
4.

Continue normal operation.

If failure indications remain on:
4.

Land as soon as practical.

If CADC caution light goes off:

FLCS Electronic Malfunctions

2.

Only two single electronic failures are reported inflight. These are a branch power supply failure (BRKPWR DEGR PFL) and a coil current monitor (CCM)
detected failure (FLCS CCM FAIL PFL). All othersingle electronic failures (rate gyro and accelerometer
failures, stick sensor failures, single axis failures, andsingle branch failures) are reported 2 minutes after
WOW as an FLCS SNGL FAIL PFL. A momentary,mild transient might be felt when a single electronic
failure occurs. Even though display of a single failure
is not always provided in flight, a single FLCS failure
can be reset.

Check for an ENG MACH FAIL PFL.

If ENG MACH FAIL PFL is still present:
3.

Continue flight and observe throttle limitation, if supersonic. Refer to PILOT FAULT
LIST - ENGINE, this section.

If CADC caution light remains on:
2.

AOA-Crosscheck with airspeed.
Use AOA indications with caution.

3110

T.O. GR1F16CJ1
If a branch fails due to a branch power supply
problem, BRK PWR DEGR PFL is displayed. If power
fails in branch A, B, or C, then one AOA and one air
data input are also failed. This condition is displayedby FLCS AOA FAIL and FLCS ADC FAIL PFL's inaddition to the BRK PWR DEGR PFL. A single
branch power supply failure may also cause FLCS
dual fail indications if the ADV MODE switch is notin the depressed position. This dual failure can bereset as long as the ADV MODE switch is in the
depressed position prior to reset. Anytime BRK PWR
DEGR occurs, FLCS power supply status should bechecked by placing the FLCS PWR TEST switch to
the TEST position. If any one of the four FLCS PWR
indicator lights does not illuminate, then that branch
is not powered and the indicated toe brake in theindicated channel is inoperative.
If the CCM detects an erroneous output command fromone of the three servo amplifiers which drive the three
ISA servo valves, a backup servo amplifier is energized.
FLCS CCM FAIL PFL is then displayed. An FLCS reset
may bring the tripped amplifier back on line.
If a dual electronic failure occurs (i.e., two pitch rate
gyros, two roll stick sensors, or two power supplies),
the FLCS warning light illuminates and an FLCS
DUAL FAIL PFL occurs. A reset of a dual failure may
be attempted. However, reset is only possible back tothe single failure condition, since the single failure islatched if it still existed at the time of the second
failure. Branch power status should again be checkedwith the FLCS PWR TEST switch.
There is a remote possibility that a series of FLCS
failures during maneuvering on the CAT I/III AOA
limiter can result in centering of and lack of response
from one or both horizontal tails. The series of failures
starts with a FLCS CCM FAIL PFL and can quickly
progress to ISA FAIL PFL's and a FLCS dual failure.
Centering of and loss of horizontal tail response only
occurs in conjunction with FLCS dual electronic failindications. This should not be confused with a loss ofcontrol associated with a departure from controlled
flight. If the aircraft suddenly becomes nonrespon
sive in pitch and FLCS DUAL FAIL indications arepresent, perform a FLCS reset.
It is possible to receive FLCS dual fail indications
following a single failure of the ADV MODE switch
when attempting/during ATF operations. The dual
redundant switch has two outputs, each of which are
split for use by the fourchannel FLCS. If one side of
the switch fails, the FLCS warning light illuminates
and an FLCS DUAL FAIL PFL occurs. This dual
failure can be reset as long as the ADV MODE switch
is disengaged prior to reset.
LESS b1q Vibrations associated with gun firing may
cause either a single failure or a dual failure of a

FLCS accelerometer to be declared. These failure
declarations occur during gun firing because the four
outputs of the accelerometer do not always compare
within present software tolerances.
FLCS SINGLE ELECTRONIC FAILURE
If BRK PWR DEGR or FLCS CCM FAIL PFL occurs:
1.

Establish 1g flight and airspeed less than 400
knots (subsonic).

2.

FLCS RESET switch-RESET.

If failure indications go off:3.

Continue normal operation.

If failure indications remain on:
3.

FLCS PWR TEST switch-TEST.
Observe FLCS PWR lights and determine
brake and brake channel affected. If branch
A, B, or C FLCS PWR light fails to illuminate,
use a maximum of 11 degrees AOA for
approach, landing, and twopoint aero
dynamic braking.

4.

BRAKES channel switch-Change channels(if required).

5.

Land as soon as practical.

FLCS DUAL ELECTRONIC FAILURE
If FLCS DUAL FAIL PFL occurs:
1.

Establish 1g flight and airspeed less than 400
knots (subsonic).

2.

ADV MODE switch-Depress.
The ATF NOT ENGAGED caution light may
illuminate shortly after depressing the ADV
MODE switch.

3.

FLCS RESET switch-RESET.
Reset may clear the FLCS warning light;
however, the single failure is still present.

If FLCS warning light goes off and no FLCS PFL's arepresent:
4.

Continue normal operation, but do not use
ADV MODE switch.

If FLCS warning light goes off and an FLCS PFL isstill present:
4.

FLCS PWR TEST switch-TEST.
Observe FLCS PWR lights and determine
brake and brake channel affected. If branch
A, B, or C FLCS PWR light fails to illuminate,
use a maximum of 11 degrees AOA for
approach, landing, and twopoint aero
dynamic braking.
Change 13111

T.O. GR1F16CJ1
5.

BRAKES channel switch-Change channels(if required).

6.

Land as soon as practical.

NOTE
External fuel cannot be transferred in
OFF or RAM. Consider jettisoningtanks to decrease drag if range iscritical and ECS cannot be turned on
for short periods of time to transfer
fuel.

If FLCS warning light remains on:
4.

FLCS PWR TEST switch-TEST.
Observe FLCS PWR lights and determine
brake and brake channel affected. If branch
A, B, or C FLCS PWR light fails to illuminate,
use a maximum of 11 degrees AOA for
approach, landing, and twopoint aero
dynamic braking.

If failure indications go off:
It may take up to 15 minutes for ramair cooling to
extinguish the light.

5.

BRAKES channel switch-Change channels(if required).

If failure indications remain on:

6.

Land as soon as possible.
NOTE
FNo significant flying qualities degradation should occur; however, with an
FLCS dual failure, the FLCS has no
redundancy.

FTwo minutes after WOW, the FLCS

FAULT caution light illuminates andan FLCS SNGL FAIL PFL occurs.

FLCS Temperature Malfunction

4.

5.

Land as soon as practical.

Land as soon as possible.

Autopilot Malfunctions
If the autopilot fails to receive needed data or AOA
exceeds 15 degrees, the FLCS FAULT caution lightilluminates, the FLCS A/P FAIL PFL is displayed,
and the autopilot automatically disengages. If
engagement is attempted with a failure or degradation, or AOA above 15 degrees the autopilot does notengage and FLCS A/P FAIL PFL occurs. If ATF is
engaged or if engagement is attempted, this failure
also causes the ATF NOT ENGAGED caution light toilluminate.

If sensors in any two branches of the FLCC detect
temperatures in excess of 75_C, the FLCS FAULT
caution light illuminates and the FLCS HOT TEMPPFL is displayed.

If FLCS A/P FAIL PFL occurs:
1.

Establish 1g flight.
Below 15 degrees AOA.

If an FLCS HOT TEMP PFL occurs:

2.

FLCS RESET switch-RESET.

1.

Airspeed-400 knots maximum (subsonic).

If PFL clears:

2.

Altitude-25,000 feet MSL maximum.
If possible, descend below 15,000 feet MSL.

3.

3.

AIR SOURCE knob-RAM.

FWith the ECS shut down or the AIR

SOURCE knob in OFF or RAM, thegsuit does not inflate and PBG isdisabled.

F PX III If AIR SOURCE knob is placed to

OFF or RAM, OBOGS is inoperative.
Activate EOS if OXY LOW warning
light illuminates above 10,000 feetcockpit altitude.

3112Change 1

Continue normal operation.

If PFL remains, autopilot cannot be engaged.
The FLCS FAULT caution light and the FLCS A/P
DEGR PFL are displayed during nonTF operations
if the autopilot is engaged but is not providing theselected function.
The PFL occurs under the following conditions:

S Attitude hold modes-After 12 seconds of operation
outside of attitude limits with no stick input.

S Roll attitude hold mode-After 12 seconds ofoperation with the autopilot roll command at its
maximum authority.

T.O. GR1F16CJ1

S Pitch attitude and altitude hold modes-After 5

seconds of operation with the autopilot pitch
command at its maximum authority, no stick
inputs, and aircraft not correcting back toreferenced altitude/attitude.

3.

DIGITAL BACKUP switch-Cycle to BACKUP, then back to OFF.

If DBU ON warning light goes off:
Verify that DBU is no longer present on the FLCS
page of the MFD.

If FLCS A/P DEGR PFL occurs:

4.

FLCS RESET switch-RESET (if required).

1.

Maneuver aircraft into autopilot envelope.

5.

2.

FLCS RESET switch - RESET.

Land as soon as practical.
Do not exceed 500 knots/0.9 mach.

If DBU ON warning light remains on:

If PFL clears:

4.

DIGITAL BACKUP switch-BACKUP.

3.

5.

Airspeed-500 knots/0.9 mach maximum.
If possible, slow to 300 knots. Avoid abrupt
control inputs. Restrict bank angle changes
to less than 90 degrees.

6.

Controllability-Check.
Lower LG at safe altitude and check handling
qualities at 1113 degrees AOA. A mild
noseup transient of approximately 2 degrees
occurs if LG is lowered below 200 knots.

7.

FLCS PWR TEST switch-TEST.
Observe FLCS PWR lights and determine
status of toe brakes. If branch A, B, or C FLCSPWR light fails to illuminate, use a maximumof 11 degrees AOA for approach, landing, andtwopoint aerodynamic braking.

8.

BRAKES channel switch-Change channels (ifrequired).

9.

Land as soon as possible.
Plan a straightin approach.

Continue normal operation.

If PFL remains:
3.

Disengage autopilot.

DBU ON Warning LightDBU is entered automatically by detection of a primary
software fault. Such a fault is assumed to exist if poweris lost in three branches, if three processors fail, or if athreebranch combination of power and processor
failures occur. If automatic DBU engagement occurs,
toe brake operation may be affected.
Due to the simplicity of the DBU, fault tolerance isreduced and detailed FLCS failure information is
not provided. DBU is intended as a backup mode thatprovides a transition to cruise flight and a return tobase capability in the event that a software fault has
been encountered within the FLCC.
To transfer back to primary mode after an automatictransfer to DBU, the DIGITAL BACKUP switch must
be cycled to BACKUP and then back to OFF. Pitch
transients are possible due to horizontal tail positionchanges when the DIGITAL BACKUP switch is
cycled back to OFF. These transients can be
minimized by stabilizing at 1g prior to reselecting
OFF. If ATF was engaged prior to automatic DBU
engagement, a 3g incremental nonrolltowings levelflyup is present. Depressing the paddle switch
interrupts the flyup command. Placing the DIGITALBACKUP switch to BACKUP cancels the flyup
command.

TF FAIL Warning Light
The TF FAIL warning light illuminates when a TF
malfunction is detected by the NVP or when SWIM
failures are detected by the FLCS.
If TF FAIL warning light illuminates:
1.

Altitude-As required.
Climb to minimum enroute altitude (MEA) or
depart low altitude environment, if required.

2.

Paddle switch-Depress (if required).
This action interrupts the flyup in ATF or
manual TF (if enabled).

3.

PFLD-Check.
If a SWIM PFL is displayed, the TF
malfunction was detected by one or more
SWIM monitors.

4.

CARA, EGI/INS, and PX II FCC,
MMC-Check for proper operation.

If DBU ON warning light illuminates:
1.

Establish 1g flight.
Do not use abrupt control inputs or makerudder inputs during rolls.

2.

Airspeed-500 knots/0.9 mach maximum.
If possible, slow to 300 knots.

PX III

Change 13113

T.O. GR1F16CJ1
If SWIM ATF FAIL, SWIM NVP FAIL, SWIM RALT
FAIL, or SWIM SCP FAIL PFL is displayed:
5.

Paddle switch-Release.
NOTE
If the malfunction was detected by
SWIM and this malfunction is nolonger present, releasing the paddle
switch resets the SWIM monitors,
cancels the flyup, and extinguishes
the TF FAIL warning light.

Certain LEF malfunctions do not activate the FLCS
LEF LOCK PFL. The presence of higher than normal
buffet levels during maneuvering flight and reduced
directional stability in the high AOA region are
indications that the LEF have failed to schedule
properly.
If an FLCS LEF LOCK PFL occurs or a malfunction
is suspected (without an FLCS LEF LOCK PFL):
1.

If SWIM ATF FAIL, SWIM NVP FAIL, SWIM RALT
FAIL, or SWIM SCP FAIL PFL does not clear orrecurs:
6.

Exceeding 12 degrees AOA reduces
departure resistance. Limit rolling
maneuvers to a maximum bank angle
change of 90 degrees and avoid rapidroll rates.

Discontinue TF operations.

If SWIM ATTD FAIL or SWIM VEL FAIL PFL isdisplayed:
5.

Paddle switch-Release.

6.

Discontinue TF operations.

2.

If no SWIM PFL was present (NVP malfunction):
5.

Paddle switch-Release.

6.

Perform TFR BIT.

Discontinue TF operations.

4.

Airspeed - Decelerate to subsonic flight if
supersonic.

5.

LE FLAPS switch - LOCK (after LG is down).
Lock LEF's in landing configuration at final
approach airspeed at a safe altitude. This
makes final approach and landing as normalaspossibleandprotects
againstuncommanded LEF excursions close to theground.

6.

Land as soon as practical.
With the LEF at or near full up, there are no
unique control inputs required. A small
increase in airspeed may be noted compared
to a normal landing approach at 11 degrees
AOA. With the LEF at or near full down, theaircraft may tend to float in ground effectand a slight forward stick force may berequired.

LEF Malfunction (Symmetric)
A symmetric LEF malfunction may be indicated by
an FLCS warning light and an FLCS LEF LOCKPFL. These indicate that one or both of the LEF
branches have malfunctioned, that the asymmetry
brakes have been activated, or that the LEF havebeen manually locked.
LEF's will stop and remain fixed in position when an
FLCS LEF LOCK PFL occurs. LEF should remain
symmetrical (within 10 degrees).

3114

Continue flight.

If the FLCS warning light does not reset or amalfunction is suspected (without an FLCS LEFLOCK PFL):

If NVP malfunction still exists:
7.

FLCS RESET switch - RESET.

If FLCS warning light resets:
3.

Further TF operations should not be
attempted after the occurrence of a
SWIM ATTD FAIL or SWIM VEL FAILPFL.

AOA - 12 degrees maximum.

During engine shutdown:
7.

MAIN PWR switch - Do not place to OFF until
engine rpm has reached zero.

T.O. GR1F16CJ1heading drift will increase as AOA increases. Lower
the nose immediately after touchdown. Directional
control should not be a problem.
Placing MAIN PWR switch to OFF
before hydraulic pressure is lost may
cause damage to two LEF shafts.

If LEF asymmetry occurs:
1.

AOA - 610 degrees.

LEF Malfunction (Asymmetric)
The most likely cause of asymmetric LEF's is amechanical disconnect in one of the LEF drive trainsaccompanied by a failure of the asymmetry brake.
This failure may not activate the FLCS LEF LOCKPFL or illuminate the FLCS warning light. The firstindication of an asymmetry is an uncommanded roll.
The failed LEF may be as much as 90 degrees up or
down. Adequate roll control is available below 10
degrees AOA at subsonic speeds. Use lateral stick forroll control. Use roll trim to reduce lateral stick force
as required. Do not attempt to achieve coordinatedflight. Avoid using rudder except to reduce sideslip
when stores are jettisoned or to aid in maintaining
desired ground track during the final part of landing
approach. Do not use rudder trim. If the yaw is awayfrom the failed LEF (i.e., nose left yaw with right LEF
failed up), rudder inputs to reduce resulting sideslip
actually aggravate the situation by increasing roll
control requirements. Accepting some sideslip
reduces roll control requirements. To prevent
excessive sideslip, maintain AOA as low as practical.
Banked flight reduces the amount of heading change
due to sideslipinduced heading drift. Lock the good
LEF as close to symmetrical as possible to aid in roll
control and to prevent transients caused byautomatic scheduling. Monitor fuel consumption
since significantly higher thrust is required tocompensate for the increased drag.
Selectively jettison stores to reduce asymmetry andsideslip and reduce fuel weight as necessary to
reduce approach speed. Perform a controllability
check. The aircraft tends to roll into the wing withthe least lift (i.e., the heavy wing). If the LEF is failed
up, lift on that wing is less. If the LEF is failed down,
lift on that wing is more or less depending on thefailed LEF position and the position at which theother LEF is locked. If there is a significant
crosswind, diminish crosswind effects, if possible, bylanding with the heavy wing upwind.
Fly a shallow, straightin approach at approximately
8 degrees AOA (fly no lower than 6 degrees AOA) withminimum roundout for touchdown. Immediately
prior to touchdown, use rudder as required to alignthe aircraft with the runway. Reduce the rate ofdescent somewhat prior to touchdown, as required,but do not flare or raise the nose above 10 degrees
AOA because available roll control is reduced andFExceeding 10 degrees AOA may resultin insufficient roll authority. Limitrolling maneuvers to gentle roll in witha maximum bank angle of 30 degrees.

FFlying a fast approach (lower than 6

degrees AOA) presents additional
control difficulties caused by a change
in the path of the disturbed airflow
coming off the failed LEF.

2.

Lateral stick/roll trim - As required.

Minimize rudder inputs. Use rudder as
required to reduce sideslip when jettison
ing stores or to aid in maintaining desired
ground track during the final part oflanding approach. Do not use rudder trim.
3.

LE FLAPS switch - LOCK.
Lock operating LEF as near symmetrical as
possible.

4.

Stores - Jettison (if required).
Consider selective jettison of stores from theheavy wing as a means to reduce roll control
requirements. Refer to SELECTIVE JETTISON, this section.

5.

Fuel weight - Reduce (if feasible/required).

Reduce fuel weight if fatigue is not a
factor. Fuel flow is significantly higher
with an LEF failed full up or down andmust be considered during recovery.
6.

Controllability - Check.
Lower LG at a safe altitude and check
handling qualities at 68 degrees AOA.

7.

Land as soon as practical.
3115

T.O. GR1F16CJ1
Stick InterferenceFPrior to landing with a significant

asymmetric LEF condition, consider
aircraft configuration, pilot experiencelevel, pilot arm fatigue, airfield facilities, weather, winds, and light condi
tions (day/night). If conditions are notfavorable, a controlled ejection isrecommended.

FIf crosswind component is greater than

10 knots, choose a runway, if possible,
which allows landing with the heavy
wing upwind. Fly a shallow, straightinapproach at approximately 8 degrees
AOA (fly no lower than 6 degrees AOA)
with minimum roundout for touch
down. Use rudder, as required, to alignaircraft with the runway immediately
prior to touchdown.

8.

Stick interference can occur at anytime for a numberof reasons. Known hazards include intentional orunintentional input by the passenger/pilot not incontrol, the utility light/adjustable sliding holder, or
the right lapbelt buckle.
Contact between the right leg/knee and the stick can
result in an unintentional right roll command. The
resulting right roll may be perceived as a flightcontrol problem especially in the D aircraft when apassenger/pilot not in control unknowingly interfereswith the stick. The probability of interference isincreased with feet on the floor versus feet on therudder pedals, bulky personal equipment, or gsuit
inflation.

Stick - Lower the nose immediately after
touchdown.

Until WOW, forward stick pressure inexcess of approximately 2 pounds
results in full trailing edge down
deflection of the horizontal tails with
reduced directional control and wheel
braking effectiveness.
If departureend arrestment is required:
9.

Prior to any ejection seat movement,
clear the area around the stick.

HOOK switch - DN.

D The passenger/pilot not in control
must take care not to interfere with the
stick as a result of leg/knee movement
or gsuit inflation. If stick interferenceis suspected, the pilot in control should
depress the paddle switch to eliminate
undesired inputs.

Contact between the utility light/adjustable sliding
holder and the stick can result in unintentional stickcommands. Unintentional stick commands may beperceived as a flight control problem. If
uncommanded stick inputs are encountered, checkfor interference between the stick and utility
light/adjustable sliding holder.

Trim Malfunction
Trim malfunction is detected by an increase in stickpressure required to maintain the desired attitude orby a lack of response to stick trim inputs.
1.

TRIM/AP DISC switch - DISC, then NORM.

If normal operation is not restored:
2.

TRIM/AP DISC switch - DISC.
Autopilot cannot be engaged.

3.

ROLL and PITCH TRIM wheels - As required.

3116

Failure to properly secure utility
light/adjustable sliding holder can
result in stick interference. The
adjustable sliding holder may become
loose and come in contact with the
stick. D If stick interference isconfirmed, undesired inputs may beeliminated by placing the STICKCONTROL switch to the appropriate
position and depressing the paddle
switch.

T.O. GR1F16CJ1
If the seat is moved after an object (especially theright lapbelt buckle) becomes lodged between the seatand stick, unintended stick inputs can occur.
Reversing the direction of the initial seat movement
should correct the situation.

FThe lapbelt should remain fastened at

all times. If the lapbelt is opened inflight, caution must be taken to insure
the right lapbelt buckle does not
become lodged between the ejection
seat and stick.

FDo not move the seat with the lapbelt

disconnected.

FUEL MALFUNCTIONS
Fuel Management System PFL
An FMS FAIL PFL indicates that the fuel reference
voltage supplied to the PX II FCC, PX III MMC is outof tolerance. Fuel system effects associated with the
PFL range from degraded FCC fuel computations(e.g., BINGO fuel) to degradation/failure of the fuelquantity indicating system. If an FMS FAIL PFL
occurs, monitor the FUEL quantity indicator for
proper operation.
Fuel Leak
A fuel leak may first be noticed by visual means, fuelimbalance, an unexpected FWD or AFT FUEL LOW
caution light, or an unusually high fuel flow
indication. Monitor the totalizer to determine
whether or not a leak exists.
If a fuel leak is suspected (indicated by abnormallyhigh fuel flow, by totalizer decreasing at abnormal
rate, or by visual means):
1.

Range - Maximize.
If a suitable landing field is not within gliding
distance, consider increasing airspeed andaltitude (without the use of AB) to maximize
range by using fuel which would otherwise be
lost.

Avoid negative g flight when either
reservoir is not full.

If fuel flow is abnormally high:2.

ENG FEED knob - OFF.
Leak is in the engine feed line or engine
components.

3.

Land as soon as possible.
Consider stores jettison if range is critical.
Consider an SFO. Refer to SIMULATEDFLAMEOUT (SFO) LANDING, this section.

If fuel flow is normal:
2.

ENG FEED knob - NORM.
A NVP TFR FAIL PFL and a flyup can occur
when NORM is reselected while operating in
TFR.

If leak is from the forward system:
3.

FUEL QTY SEL knob - Out of NORM.
This action stops automatic forward fueltransfer.

If external tanks contain fuel:4.

TANK INERTING switch - TANK INERTING
to reduce internal tank pressurization.

If external tanks are not installed or when they are
empty:
5.

AIR REFUEL switch - OPEN.

6.

Land as soon as possible.
Consider stores jettison if range is critical.

If aft fuel imbalance exists (aft CG):
7.

AOA - 15 degrees maximum.

Aft fuel heavy (red portion of AL
pointer showing) results in increased
susceptibility to departure and deep
stall conditions. Limit AOA and avoid
maximum command rolling maneu
vers.

If twopoint aerodynamic braking isused with an aft CG, pitch overshoots
may
occur
and
the
nozzle,speedbrakes, and ventral fins may
contact the runway.
3117

T.O. GR1F16CJ1
Fuel Low

If external fuel has not transferred:

A fuel low caution light may be caused by a fuel leak,trapped external fuel, PX III trapped CFT fuel, a fuelimbalance between the forward and aft systems,
prolonged AB operation, or a fuel sensing problem.

6.

The FWD FUEL LOW and AFT FUEL LOW caution
lights indicate either a reservoir fuel level sensing
system malfunction or that reservoir tank quantities
are less than:

7.

C

D

FWD 400 pounds

FWD 250 pounds

AFT 250 pounds

AFT 400 pounds

PX III If CFT fuel has not completely transferred:

Fuel flow - Reduce to the minimum required tosustain flight below 6000 pph.

Go to TRAPPED CFT FUEL, this section.

If forward and aft fuselage fuel is not properly
balanced:
8.

Go to FUEL IMBALANCE, this section.

If fuel is properly balanced:
NOTE
A fuel line between the reservoir and
FFP may be ruptured, causing fuel tocycle between tanks in the same
system.

If FWD FUEL LOW and/or AFT FUEL LOW caution
light illuminates:
1.

Go to TRAPPED EXTERNAL FUEL, thissection.

9.

Land as soon as possible.

If reservoir tanks indicate full:
4.
Limit fuel flow to the minimum
required to sustain flight while thecause of the fuel low light(s) isdetermined. Avoid negative g flightwhen either reservoir is not full.
2.

3.

ENG FEED knob - NORM.
A NVP TFR FAIL PFL and a flyup can occur
when NORM is reselected while operating in
TFR.
FUEL QTY SEL knob - RSVR.
Leave FUEL QTY SEL knob out of NORM if
FUEL quantity indicator displays erroneous
information.

If either or both reservoir tanks are low:
NOTE
Fuel flow indications may fluctuatewith either reservoir empty.
4.

Land as soon as possible.
Consider an SFO. Refer to SIMULATEDFLAMEOUT (SFO) LANDING, this section.

If a fuel leak is suspected (indicated by abnormallyhigh fuel flow, by totalizer decreasing at abnormal
rate, or by visual means):
5.

Go to FUEL LEAK, this section.

3118

FUEL QTY SEL knob - TEST.

If AL and/or FR pointers test bad, or FUEL quantity
indicator is inoperative:
5.

Land as soon as possible.
Consider an SFO. Refer to SIMULATEDFLAMEOUT (SFO) LANDING, this section.

If AL and FR pointers test good:
6.

Individual fuel quantities - Check andcompare with totalizer.
Monitor reservoir tanks to insure they are
maintained full.

7.

Land as soon as practical.

Hot Fuel/Oil or Gravity Feed
Gravity feed from the reservoirs to the engine occurs
after loss of the main and standby generators andfailure of either hydraulic system A or the FFP.
Failure of the FFP may be detected by improper fuelbalance. Fuel continues to be transferred to bothreservoirs by siphoning action. Fuel distribution
cannot be manually or automatically controlledduring gravity feed. Minimize aircraft maneuvering
for duration of flight. Due to the ingestion of air into
the engine fuel system, engine flameout may occur
when either reservoir tank empties. If the standby
generator is operating, one boost pump continues totransfer fuel from the C forward, D aft reservoir
to the FFP.

T.O. GR1F16CJ1
Hot fuel, as indicated by the FUEL/OIL HOT caution
light, may result from high speed flight or fuelsystem/heat exchanger malfunctions. Excess fueltemperatures may result in engine malfunctions.
Engine flameout may occur at low flow rates
associated with the landing pattern due to hot fuel.
Fuel flow above 4000 pph minimizes fuel temperature rise.
GE 129Hot engine oil, as indicated by the
FUEL/OIL HOT caution light, may also result fromhigh speed flight or fuel/oil heat exchanger
malfunctions. Sustained engine operation withexcessive oil temperature may result in engine
damage and/or failure.

If FUEL/OIL HOT caution light illuminates orgravity feed situation exists:

1.

AIR REFUEL switch-Check CLOSE.

2.

TANK INERTING switch-Check OFF.

3.

Altitude-10,000 feet maximum (if practical).
Minimize aircraft maneuvering for durationof flight.

4.

Fuel flow-4000 pph minimum until landing
is assured when in a hot fuel situation.

If FUEL/OIL HOT caution light goes off:
5.

Land as soon as practical.

If FUEL/OIL HOT caution light remains on or
gravity feed situation exists:
5.

Land as soon as possible.
Consider an SFO. Refer to SIMULATEDFLAMEOUT (SFO) LANDING, this section.

Fuel Imbalance

FEngine flameout may occur at low fuel

Refer to figure 315. A fuel imbalance when notcarrying an external fuel tank(s) indicates a system
malfunction. A fuel imbalance when carrying an
external fuel tank(s) may be the result of normal
system operating tolerances.

flow rates when in a hot fuel situation.

FEngine flameout may occur wheneither reservoir tank empties if agravity feed condition exists.

Fuel Imbalance IndicationsC

D

FUEL IMBALANCE WARNING (AFT CG)

FUEL IMBALANCE WARNING (AFT CG)

Fuel imbalance warning (red portion of AL
pointer) shows when forward fuselage fuel(FR) indication is less than aft fuselage
fuel (AL) indication.

Fuel imbalance warning (red portion of AL
pointer) shows when forward fuselage fuel(FR) indication is more than 1350 pounds
less than aft fuselage fuel (AL) indication.

GR1F-16CJ-1-0127X37

Figure 315.

Change 13119

T.O. GR1F16CJ1NOTE
FAny correction required per total fuelLimit fuel flow to the minimum
required to sustain flight while thecause is determined. Avoid negative gflight when either reservoir is not full.

quantity usage with internal fuel only
indicates a system malfunction.

FMore than one correction per total fuelquantity usage with either a 300gal
lon fuel tank or two 370gallon fueltanks indicates a system malfunction.

If aft fuel imbalance exists (aft CG):
2.

AOA-15 degrees maximum.

FMore than two corrections per total

fuel quantity usage with either a
300gallon fuel tank and two 370gal
lon fuel tanks or two 600gallon fueltanks indicate a system malfunction.

FMore than three corrections per total

fuel quantity usage with a 300gallon
fuel tank and two 600gallon fuel tanks
indicate a system malfunction.

F PX III Placing the ENG FEED knob toeither FWD or AFT during externaltank fuel transfer may cause some fuelto enter empty CFT's.

A fuel imbalance is indicated by the red portion of the
AL pointer. This may be caused by an FFP
malfunction, fuel leak, uneven or partial refueling(either ground or AR), or a malfunction of theautomatic forward fuel transfer system. A fuelimbalance may also occur as a result of a main
generator failure and an inoperative FFP (FFP
malfunction or system A hydraulic failure). In thiscase, boost pump No. 3 still transfers fuel from the C
forward, D aft reservoir to the engine, creating theimbalance.
An unexpected FWD or AFT FUEL LOW caution
light may also be an indication of fuel imbalance;
however, verify that forward fuselage fuel and aftfuselage fuel (as indicated by AL and FR pointerswith FUEL QTY SEL knob in NORM) are notproperly balanced and that a leak does not exist
before selecting FWD or AFT ENG FEED.
If fuel imbalance is indicated by AL and FR pointerswith FUEL QTY SEL knob in NORM:
1.

Fuel flow-Reduce to the minimum required
to sustain flight below 6000 pph.

3120

Aft fuel heavy (red portion of AL pointershowing) results in increased suscepti
bility to departure and deep stallconditions. Limit AOA and avoid
maximum command rolling maneuvers.
If a fuel leak is suspected (indicated by abnormallyhigh fuel flow, by totalizer decreasing at abnormal
rate, or by visual means):
3.

Go to FUEL LEAK, this section.

If a fuel leak is not suspected:
4.

Fuel quantities-Check.
Use the FUEL QTY SEL knob to determine if
a trapped fuel condition exists. Refer to
TRAPPED EXTERNAL FUEL, this section, if
required.

5.

ENG FEED knob-FWD or AFT.
Use only to correct a forward and aft fuselage
fuel imbalance and not to correct imbalances
between reservoirs. Do not exceed 25,000 pph
fuel flow while balancing fuel.

If imbalance is not corrected:
6.

Land as soon as practical.

If twopoint aerodynamic braking is usedwith an aft CG, pitch overshoots may
occur and the nozzle, speedbrakes, andventral fins may contact the runway.
If proper distribution is attained:
6.

ENG FEED knob-NORM.
A NVP TFR FAIL PFL and a flyup can occur
when NORM is reselected while operating in
TFR.

7.

Fuel balance-Monitor.

T.O. GR1F16CJ1external wing tanks, the centerline
tank, or all three external tanks. Witha three tank configuration, the firstindication that the centerline tank isfeeding is after the external wingtanks are emptied.

Trapped External FuelFA TRP FUEL indication in the HUD

may be a symptom of an external fuelleak. If a fuel leak is suspected(indicated by abnormally high fuelflow, by totalizer decreasing at abnor
mal rate, or by visual means), refer to
FUEL LEAK, this section.

7.

ENG FEED knob-NORM.
A NVP TFR FAIL PFL and a flyup can occur
when NORM is reselected while operating in
TFR.

8.

Stick-Pulse aircraft in pitch several times byapplying differential g forces of approximately"2g.

FWith trapped external fuel, the total

izer does not indicate total usable fuel.
Usable fuel is the totalizer quantity
less the external fuel quantity.
NOTE

PX III If either INT WING & CFT
indication is greater than 700 pounds
and an external tank is empty, go to
TRAPPED CFT FUEL, this section.

Certain malfunctions can cause fuel to be trapped in
the external tank(s). The tank(s) in which fuel istrapped can be detected by a periodic check of
external tank fuel quantities.
Accomplish steps 1 through 8 and 9 (if required)without delay:
NOTE

If the AIR REFUEL switch was initially found in
CLOSE (step 2), perform step 9. If the AIR REFUEL
switch was initially found in OPEN (step 2), omit
step 9.
9.

AIR REFUEL switch-OPEN (1 second), then
CLOSE.
Open or close AR door at or below 400
knots/0.85 mach.

10. External tank fuel quantity-Monitor.
The time required to observe fuel transfer if
the malfunction is corrected can vary from13 minutes (for a full centerline tank) to1012 minutes (for three external tanks with
500 pounds fuel in each) if reservoir tanks
are full (i.e., both air ejectors are off).

Repeating or undoing any steps may
delay transfer.
1.

Fuel flow-Minimize.

2.

AIR REFUEL switch-Confirm in CLOSE.

3.

AIR SOURCE knob-Confirm in NORM or
DUMP.

4.

TEMP knob-MAN and adjust for comfort.
This action usually increases ECS airpressure for external fuel transfer.

5.

TANK INERTING switch-TANK INERTING to reduce internal tank pressurization.

6.

EXT FUEL TRANS switch-WING FIRST.

If a trapped external fuel condition is
not discovered until either reservoir
tank is less than full or a fuel low lightis on, sufficient fuel transfer from theexternal tank(s) may not occur even if
the malfunction is corrected. Consider
fuselage fuel to be the only usable fuel.
NOTE
If trapped external fuel occurs after air
refueling and completion of checklist
steps did not correct the malfunction,
consider descending well below thefreezing level to unfreeze the externalpressurization and vent valve. Cycling
the AR door at lower altitude may
restore normal operation.

NOTE
Selecting WING FIRST bypasses electrical components that, if malfunctioning, can prevent fuel transfer from

11.

Stores-Jettison (if required).
3121

T.O. GR1F16CJ1
Trapped CFT Fuel PX III

If FWD FUEL LOW and AFT FUEL LOW caution
lights are off:4.

FWith trapped CFT fuel, the totalizer

does not indicate total usable fuel.
Until fuel transfer can be established,fuselage fuel is the only available
usable fuel.

NOTE
Opening the AR door depressurizes
external tanks and removes the cause oftrapped CFT fuel. It may take from 1
minute (centerline tank) to 4 minutes(centerline tank and two 600gallon fueltanks) for external tank air pressure to
decrease to zero. With the air source
removed, CFT fuel can be transferred.
The wing turbine pump capability
limits the transfer rate of CFT fuel fromthe internal wings to the fuselage.

FIf a trapped CFT fuel condition is notdiscovered until either reservoir tankis less than full or a fuel low light is on,
sufficient fuel transfer from the CFT
may not occur even if the malfunction
is corrected. Consider fuselage fuel tobe the only usable fuel.

If an external tank transfer valve does not close
when the tank empties, air pressure is routed to the
CFT. This air may completely trap CFT fuel or may
cause as much as 800 pounds per side of CFT fuel tovent overboard and trap the remaining CFT fuel.

AIR REFUEL switch-OPEN.
Open or close AR door at or below 400
knots/0.85 mach.

5.

Typically a centerline tank transfer valve failure willaffect both left and right CFT's and an external wingtank transfer valve failure will affect the CFT on thesame side. A centerline tank failure may trap fuel inboth CFT's and both external wing tanks; in thiscondition, the CFT fuel must be transferred before
the external wing tank fuel can be transferred.

Fuel quantities-Monitor.
The time required to observe fuel transfer canvary from 1025 minutes after AR door is
opened. Because CFT fuel is combined withinternal wing fuel, the INT WING & CFT
quantity will not immediately decrease. As fueltransfers, the INT WING & CFT quantity
indication may be very erratic with jumps of
200 pounds. CFT fuel transfer is best
determined by observing a reduction in
fuselage fuel usage or an increase in fuselage
fuel.

1.

Fuel flow-Minimize.

2.

EXT FUEL TRANS switch-CFT FIRST / NO
FILL.

NOTE

3.

FUEL QTY SEL knob-Check all positions.
If INT WING & CFT quantity remains
greater than 700 pounds and fuselage fuel isdecreasing and an external tank is empty,
fuel is trapped in the CFT.

If no fuel transfer is apparent after 10
minutes with AR door open, consider
descending. A descent of 1/3 of thealtitude available may speed up theprocess by increasing air pressure
behind the CFT fuel.

If FWD FUEL LOW and/or AFT FUEL LOW caution
light is on:
4.

Stores-Retain any external tank containing
fuel; jettison any empty external tank andother stores. Refer to JETTISON, this section.

5.

AIR REFUEL switch-OPEN for 1 minutethen CLOSE.
Open or close AR door at or below 400
knots/0.85 mach.
NOTE
Opening the AR door for 1 minute vents
pressure that may prevent transfer of
CFT fuel.

3122Change 1

When each INT WING & CFT quantity is less than
200 pounds:
6.

AIR REFUEL switch-CLOSE.
NOTE
Closing the AR door repressurizes theexternal tank(s). Repressurization
may be slow because of the failed
external tank and may not be suffi
cient to obtain normal external tanktransfer rate. With EXT TANK TRANS
switch in CFT FIRST/NO FILL, the
CFT's will remain empty.

T.O. GR1F16CJ1
If fuselage fuel is not sufficient to recover theaircraft:tank is removed, the AR door can beclosed so that the fuel system pressure
will increase and improve the CFTtransfer rate. Fuel in any remaining
external tank(s) may also transfer tofill the internal wing.

NOTE
Jettison of the failed empty externaltank will immediately remove thesource of air trapping the CFT fuel.
Jettison does not improve the fueltransfer rate. However, once the failed

7.

Jettison empty external tank(s).

8.

AIR REFUEL switch-CLOSE.

Change 13122.1/(3122.2 blank)

T.O. GR1F16CJ1HYDRAULIC MALFUNCTIONS
A hydraulic system failure is indicated by illumina
tion of the HYD/OIL PRESS warning light, FLCS
FAULT caution light, and ISA ALL FAIL PFL. The
HYD/OIL PRESS warning light illuminates when
ever either hydraulic system pressure drops below
1000 psi. The ISA ALL FAIL PFL may occur prior to
the HYD/OIL PRESS warning light.
With system B hydraulic failure, perform alternate
LG extension with the LG handle up. This action
reduces the possibility of failing to unlock a LG door
actuator or the NLG extend/retract actuator due tolow or fluctuating system B hydraulic pressure.

prepare for possible LG up landing. Drag chuteoperation is normal using drag chute accumulator
pressure. Braking is available using brake/JFS
accumulator pressure. The fully charged brake/JFS
accumulators contain sufficient fluid for at least 75
seconds of continuous brake application. Use
aerodynamic braking to the maximum extent
possible. A single moderate and steady brakeapplication without cycling the antiskid should thenbe applied. After stopping, engage the parking
brake. If there is reason to believe that thebrake/JFS accumulators are depleted or that
directional control may be a problem, an approachend arrestment should be considered.
1.

Land as soon as practical.
Make smooth control inputs and plan to fly astraightin approach.

2.

ALT GEAR handle-Pull (190 knots maxi
mum, if practical).
Alternate LG extension can be used up to 300
knots; however, the NLG may not fully extend
until 190 knots. Time above 190 knots should
be minimized in case there is a leak in thepneumatic lines.

Single Hydraulic Failure

If hydraulic failure is due to structural
damage (e.g., battle damage, midair
collision, bird strike, fire, or hardlanding), the other system may bedamaged and failure can occur withlittle warning. The HYD PRESS
indicator may show normal pressure
until system fluid is depleted.

F NWS is not available following alter
nate LG extension.

SYSTEM A FAILURE

F Do not depress the ALT GEAR reset

The FLCS ISA's are operating in the nonredundantmode and the speedbrakes and FFP are inoperative.
1.

Land as soon as practical.
Make smooth control inputs and plan to fly astraightin approach.

2.

System B HYD PRESS indicator-Monitor.

3.

Fuel balance-Monitor.
Fuel distribution must be controlled manually.

button while pulling the ALT GEAR
handle. This action may preclude
successful LG extension.

3.

If LG handle does not lower, selectBRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND. PW 229
Nozzle remains closed resulting inhigher than normal landing thrust.

SYSTEM B FAILURENOTEEPU RUN light on may indicate a dual
hydraulic or PTO shaft failure.
The FLCS ISA's are operating in the nonredundantmode and normal braking, NWS, AR door operation,
gun operation, and normal LG extension are lost. Low
hydraulic pressure may cause one or more LG
actuators to remain locked in the LG up position. LG
extension should be attempted in sufficient time to

LG handle-DN. (Use DN LOCK REL button
if required.)

4.

HOOK switch-DN (if required).
Braking is available using brake/JFS accu
mulators only. To avoid brake activation andloss of brake/JFS accumulator pressure, do
not rest feet on brake pedals. If the brake/JFS
accumulators are depleted or if directional
control may be a problem, consider anapproachend arrestment. Refer to CABLE
ARRESTMENT, this section.
3123

T.O. GR1F16CJ1
After landing:
5.

NOTE
Before landing, confirm that the EPU
operates (EPU run light is on) with thethrottle in IDLE. If the EPU run lightgoes off, refer to ABNORMAL EPU
OPERATION, this section.

Stop straight ahead and engage parking brake.

FBrakes should be applied in a single,

moderate, and steady application
without cycling the antiskid.

4.

Land as soon as possible.
Make smooth control inputs and plan to fly astraightin approach.

5.

ALT GEAR handle-Pull (190 knots maxi
mum, if practical).
Alternate LG extension can be used up to 300
knots; however, the NLG may not fully extend
until 190 knots. Time above 190 knots should
be minimized in case there is a leak in thepneumatic lines.

FBrake pedal deflection of 1/16 inch

activates the brakes and bleeds thebrake/JFS accumulators. To avoid
brake activation and loss of brake/JFS
accumulator pressure, do not rest feeton brake pedals.

FDo not attempt to taxi clear of the

runway. Loss of brake/JFS accumula
tor pressure results in the inability tostop or steer the aircraft.

F NWS is not available following alter
nate LG extension.
F Do not depress the ALT GEAR reset

Dual Hydraulic Failure
A dual hydraulic system failure can be detected bysluggishness or lack of response to flight control
inputs, decreasing pressure readings on both HYD
PRESS indicators, and associated warning andcaution lights. The EPU automatically provides
hydraulic pressure for system A when pressure ofboth hydraulic systems drops below 1000 psi. The
systems affected by dual hydraulic system failure
after the EPU is running are the same as thoseaffected by system B failure. Refer to SYSTEM B
FAILURE, this section.
1.

EPU switch-ON (if EPU run light is off).

2.

System A HYD PRESS indicator-Checkpressure increasing.

button while pulling the ALT GEAR
handle. This action may preclude
successful LG extension.

6.

If LG handle does not lower, selectBRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND. PW 229
Nozzle remains closed resulting inhigher than normal landing thrust.
7.

If hydraulic pressure does not increase or control
response is lost:3.

Eject.

If system A hydraulic pressure is restored:
3.

EPU run light-Check light on at idle thrust.

3124

LG handle-DN. (Use DN LOCK REL button
if required.)

HOOK switch-DN (if required).
Braking is available using brake/JFS
accumulators only. To avoid brake activationand loss of brake/JFS accumulator pressure,
do not rest feet on brake pedals. If thebrake/JFS accumulators are depleted or if
directional control may be a problem,
consider an approachend arrestment. Refer
to CABLE ARRESTMENT, this section.

After landing:8.

Stop straight ahead and engage parking brake.

T.O. GR1F16CJ1FBrakes should be applied in a single,

PX II If only AUX flag is in view, pitch
and roll attitude information is likely
to be erroneous due to INS autorestartin the attitude mode when other thanstraight and level, unaccelerated
flight conditions existed.

moderate, and steady application
without cycling the antiskid.

FBrake pedal deflection of 1/16 inch

activates the brakes and bleeds thebrake/JFS accumulators. To avoid
brake activation and loss of brake/JFS
accumulator pressure, do not rest feeton brake pedals.

FDo not attempt to taxi clear of the5.

Fuel balance-Monitor.

6.

EPU run light-Check light on at idle thrust.

runway. Loss of brake/JFS accumula
tor pressure results in the inability tostop or steer the aircraft.

9.

NOTE
Before landing, confirm that the EPU
operates (EPU run light is on) with thethrottle in IDLE. If the EPU run lightgoes off, refer to ABNORMAL EPU
OPERATION, this section.

Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.

System B and Generator Failure (PTO Shaft)
A PTO shaft failure is indicated by failure of
hydraulic system B, the main and standby
generators, and the FLCS PMG. The EPU should
start automatically to provide emergency hydraulicand electrical power. After accomplishing theappropriate emergency procedures, refer to EMERGENCY POWER DISTRIBUTION, this section, todetermine inoperative equipment.

7.

Land as soon as possible.
Make smooth control inputs and plan to fly astraightin approach.

8.

ALT GEAR handle-Pull (190 knots maxi
mum, if practical).
Alternate LG extension can be used up to 300
knots; however, the NLG may not fully extend
until 190 knots. Time above 190 knots should
be minimized in case there is a leak in thepneumatic lines.

NOTE
PX II The VHF radio is not powered
when the EPU is running or the main
generator is off line.

1.

EPU switch-ON (if EPU run light is off).

FNWS is not available following alter

nate LG extension.

If EPU run light is off and control response is lost:2.

FDo not depress the ALT GEAR reset

Eject.

button while pulling the ALT GEAR
handle. This action may preclude
successful LG extension.

If EPU run light is on:
3.

Throttle-As required.
9.

Stall protection may be lost. Do notretard throttle below MIL until sub
sonic.
4.

ADI-Check for presence of OFF and/or AUX
warning flags.
If warning flag(s) is in view, refer to EGIFAILURE/TOTAL INS FAILURE, this section.

LG handle-DN. (Use DN LOCK REL button
if required.)

If LG handle does not lower, selectBRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND. PW 229
Nozzle remains closed, resulting inhigher than normal landing thrust.
3125

T.O. GR1F16CJ1
10. HOOK switch-DN (if required).
Braking is available using brake/JFS
accumulators only. To avoid brake activationand loss of brake/JFS accumulator pressure,
do not rest feet on brake pedals. If thebrake/JFS accumulators are depleted or if
directional control may be a problem,
consider an approachend arrestment. Refer
to CABLE ARRESTMENT, this section.

If the autopilot is engaged when an EGI failure occurs
or during an inflight alignment (IFA) or attitudealignment (ATT), uncommanded pitch and roll flightcontrol inputs may occur.

After landing:11.

Stop straight ahead and engage parking brake.

The autopilot does not automaticallydisengage with EGI failures. Failure
to manually disconnect the autopilot
may result in an unusual aircraft
attitude and disorientation.

FBrakes should be applied in a single,

moderate, and steady application
without cycling the antiskid.

FBrake pedal deflection of 1/16 inch

activates the brake/JFS accumulators.
To avoid brake activation and loss ofaccumulator fluid, do not rest feet on
the brake pedals.

FDo not attempt to taxi clear of the

runway. Loss of brake/JFS accumula
tor pressure results in the inability tostop or steer the aircraft.

The most likely cause of an EGI failure is loss ofprimary and backup power to the EGI, causing a
shutdown. The EGI may also fail or shutdown due tointernal failures. The primary indication of an EGI
failure is the loss of attitude information (OFF and
AUX flags may be displayed on the ADI) and HSI
data. With ADI OFF and AUX flags displayed or
during an HSI failure, EGI MFL/PFL's may bedisplayed.
NOTE

12. EPU switch-OFF.
13. Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.
EGI FAILURE PX III
Most EGI failures are apparent and affect flight
operations significantly.

It is possible for the displayed ADI and/or
HUD attitude to be in error with no ADI
OFF or AUX warning flags in view andwithout an EGI or HUD MFL/PFL.
Displayed HSI and/or HUD headings
may also be in error with no HSI OFF or
ADI AUX warning flags in view andwithout an EGI or HUD MFL/PFL.
Momentary warning flags may indicate
impending failure. To detect thesefailures and maintain proper flightorientation, basic and backup instru
ments must be crosschecked.
3126

The EGI may continue to provide usefulattitude and navigation data with theoccurrence of EGI MFL/PFL's. If the
ADI OFF and AUX flags are notdisplayed and data on the ADI and HSI
can be confirmed as valid, selection of
IFA or ATT will result in loss of theexisting attitude and navigation data
and the fault condition may prevent
successful completion of the alignment.
The emergency procedures for an EGI failure withADI OFF and AUX flags in view require either aninflight or an ATT alignment. An inflight alignmentis preferred as it will provide complete navigation
capability (heading, attitude, steerpoint steering). In
addition, an inflight alignment uses GPS information and aircraft movement to derive an attitudereferenced to the horizon.
The attitude information derived from an ATTalignment is based on the aircraft attitude thatexisted when the ATT alignment was accomplished.

T.O. GR1F16CJ1The ATT alignment can provide attitude informationin approximately 10 seconds whereas an inflightalignment requires approximately 1 minute to obtain
attitude information.
There are two inflight alignment selections - AUTO
IFA and MAN IFA. AUTO IFA is a GPSbasedalignment internal to the EGI. MAN IFA is a masternavigation filter based alignment that can use GPSdata or manual fix taking if GPS is unavailable. The
four options (listed from best to worst performance)
are AUTO IFA, MAN IFA (with GPS), MAN IFA (with
fix taking), and ATT. Refer to T.O. GR1F16CJ3411
for more detailed inflight alignment procedures.

11.

12. Attitude-Maintain straight, level, and unac
celerated flight until ALIGN replaces STBY in
the HUD and ADI AUX flag is out of view.
13. Inflight alignment-Accomplish.
NOTE
FFix taking procedures may be required

as indicated on the DED MAN INFLTALIGN page.

FConstant altitude ($ 200 feet) coordi

If ADI OFF and AUX flags are in view or attitude iserroneous:
1.

EGI knob-OFF for 10 seconds.

2.

Attitude-Establish straight, level, and unac
celerated flight.

3.

EGI knob-AUTO IFA.

4.

Attitude-Maintain straight, level, and unac
celerated flight until ALIGN replaces STBY in
the HUD and ADI AUX flag is out of view.

5.

Inflight alignment-Accomplish.
NOTE
Constant altitude ($ 200 feet) coordi
nated turns (bank angle less than 45
degrees) to change heading by 45 to 90
degrees and holding the heading for 1
minute will assist completion of thealignment.

6.

7.

EGI knob-NAV after Maxg replaces ALIGN in
the HUD and RDY is removed from the DED EGI
page.
ADI, HUD, and HSI-Verify accuracy of
attitude and navigation data.

If the AUTO IFA fails to complete after 10 minutes,
consider attempting a MAN IFA with GPS or with fix
taking:8.

EGI knob-OFF for 10 seconds.

9.

Attitude-Establish straight, level, and unac
celerated flight.

10. EGI knob-MAN IFA.

Enter best available magnetic heading on the
DED MAN INFLT ALIGN page.

nated turns (bank angle less than 45
degrees) to change heading by 45 to 90
degrees and holding the heading for 1
minute will assist completion of thealignment.

14. EGI knob-NAV after Maxg replaces ALIGN
in the HUD and RDY is removed from the DED
EGI page.
15. ADI, HUD, and HSI-Verify accuracy of
attitude and navigation data.
If the MAN IFA fails to complete after 10 minutes, theattitude mode should be attempted:
16. EGI knob-OFF for 10 seconds.
17. Attitude-Establish straight, level, and unac
celerated flight.
18. EGI knob-ATT.
19. Attitude-Maintain straight, level, andunaccelerated flight until ADI OFF warning
flag goes out of view after approximately 10
seconds.
20. ADI and HUD-Verify attitude information iscorrect.
21.

C DF INSTR HDG knob-Slew HSI to match
best available magnetic heading.

INS FAILURES PX II
Most INS failures are apparent and affect flight
operations significantly.
An FLCS A/P FAIL PFL occurs if the autopilot is
engaged during an inflight alignment. The PFL can
be cleared after a full alignment is achieved by doingan FLCS reset.

3127

T.O. GR1F16CJ1
Total INS Failure PX II
A total INS failure is normally indicated by the INS
BUS FAIL PFL. The ADI freezes with OFF and AUX
warning flags in view; the HSI compass card andbearing pointer freeze; the HUD pitch ladder,heading scale, roll scale, and FPM also blank;magnetic heading is displayed only on the magnetic
compass; the FLCC AOS feedback function isdeactivated; and attitude reference is available only
on the SAI. When the INSTR MODE knob is in TCN
the HSI course deviation indicator, RANGE indicator,
and TOFROM indicator are operative, and thecapability to fly an inbound (or outbound) radial to (orfrom) a TACAN station is available.

It is possible for the displayed ADI
and/or HUD attitude to be in error withno ADI OFF or AUX warning flags inview and without an INS or HUD PFL.
Displayed HSI and/or HUD headings
may also be in error with no HSI OFF
or ADI AUX warning flags in view andwithout an INS or HUD PFL. Momentary warning flags may indicate
impending failure. To detect thesefailures and maintain proper flightorientation, basic and backup instru
ments must be crosschecked. Refer toT.O. GR1F16CJ3411 for more de
tailed inflight alignment procedures.
1.

INS knob-OFF for 10 seconds.

2.

Attitude-Straight, level, and unaccelerated.

3.

INS knob-IN FLT ALIGN.

4.

Magnetic heading - Enter.

5.

Attitude-Straight, level, and unaccelerated
until ADI OFF warning flag goes out of view
after approximately 10 seconds.

6.

Auto or manual inflight alignment-Accom
plish.
Minimum performance is available withreturn of the HUD FPM; return of MAX G
indicates full performance. INS knob canremain in IN FLT ALIGN to insure thehighest performance by continuing the INS
updating process.

3128

NOTE
Limit vertical maneuvering until theFPM is displayed on the HUD. Failure to
do so could delay or prevent completionof the inflight alignment.
If ADI OFF and/or AUX warning flag remains in view,alignment is not possible and the attitude mode
should be attempted:
7.

INS knob-OFF for 15 seconds.

8.

INS knob - ATT.

9.

Attitude - Straight, level, and unaccelerated
until ADI OFF warning flag goes out of view
after approximately 10 seconds.

10. ADI and HUD- Verify attitude information
correct.
11.

C DF INSTR HDG knob-Slew HSI to match
best available magnetic heading.

OUTOFCONTROL RECOVERY
Refer to OUTOFCONTROL CHARACTERISTICS,
Section VI, for a detailed discussion of flightcharacteristics and indications during departures,deep stalls, spins, and recoveries.
In order to prevent a departure, immediately initiate
recovery after the low speed warning tone comes on.
Prompt recovery is even more critical withheavyweight loadings or during hard maneuvering
since the airspeed bleedoff is more rapid. Additional
ly, a yaw departure may be prevented if controls arepromptly neutralized following an uncommanded
nose slice or roll hesitation.
In order to minimize time and altitude loss following
a departure, immediately release the stick andrudder controls. The aircraft should be allowed theopportunity to selfrecover. Selfrecoveries usuallyoccur within the first two postdeparture pitch
oscillations, and may take up to 1020 seconds.
Recovery is indicated by the nose dropping and the
AOA remaining below 25 degrees. Fly the aircraft at
a low AOA until airspeed reaches 200 knots or more
(if altitude permits) and recover from the resulting
dive. If the departure does not result in selfrecovery,
then the aircraft is in a deep stall or spin.
PW 229 Departures at high altitude may result in anengine stall. If in AB during an outofcontrol
situation, retard the throttle to MIL. If at MIL or
below, do not move the throttle. Do not advance thethrottle until beginning the dive recovery.

T.O. GR1F16CJ1
GE 129 Departures at high altitude may result in anengine stall. Prolonged negative g flight at a highengine thrust level may result in an engine bearing
failure. Retard the throttle to IDLE. Do not advancethe throttle until beginning the dive recovery.

Upright deep stalls may be very stable with little orno pitch motions or may be very oscillatory withlarge pitch, roll, and yaw motions. Generally, a clean
configuration results in a deep stall with a nearwingslevel pitching motion.
During upright deep stalls with a centerline store,
particularly a 300gallon fuel tank, the aircraft tends
to roll and yaw right while pitching up, and roll andyaw left while pitching down. During deep stalls with
370gallon fuel tanks, the aircraft nose motion
appears triangular. This motion is characterized by aroll and yaw right while pitching up, followed by apitch down, a hesitation, and yaw to the left.
In an upright deep stall or spin, the yaw rate limiter
automatically provides antispin controls and therudder authority limiter prevents pilot yaw
commands. The yaw rate limiter is effective inpreventing spins with almost all CAT I loadings.
However, following a yaw departure above 25,000
feet, aircraft with CAT I loadings that have all thefollowing characteristics may spin:

S Centerline store.
S Inlet mounted pod(s).
S Lateral asymmetry greater than 300 pounds at
stations 1, 2, or 3.

Upright spins following a yaw departure can bedisorienting. The initial portion of the spin ischaracterized by highly oscillatory motions and ahigh yaw rate (70 to 100 degrees per second). Initially,
the aircraft spins roughly around the aircraft's flightpath at departure. As the spin continues, the rotation
axis eventually becomes vertical. Very noticeable
forward g (eyeballs out) and sideforces are present.
In a spin, the yaw rate must be allowed to subside
before the aircraft can be recovered. This may require
20 to 30 seconds. Pitch, roll, and yaw oscillations
associated with a deep stall should not be confusedwith the continuous yaw rotation of a spin. When theyaw rotation subsides, the aircraft will either recoveror will settle into an upright deep stall.
In an inverted deep stall or spin, the yaw rate limiter
automatically provides rudder against the yaw rate.
Roll and rudder commands should be avoided. Pilot rolland rudder commands are inhibited when MPO is
engaged.

The aircraft must be rocked out of a deep stall with
the MPO switch held in OVRD until recovery iscomplete. The MPO switch allows the pilot to use thehorizontal tail surfaces to reinforce pitch oscillations
until the pitch rates are sufficient for recovery. Whensufficient nosedown pitch rate is generated to reduce
the AOA below the deep stall AOA, the aircraft will
recover.
The MPO switch must remain in the OVRD position
during pitch rocking. If the MPO switch is released,
the horizontal tails reposition to reduce AOA and may
negate any pitch oscillations. Additionally, if the
MPO switch is positioned to OVRD without any stickcommands, the horizontal tails streamline andprevent recovery.
In an upright deep stall, begin pitch rocking inphasewith nose movement; i.e., if the nose is pitching up,
pull back on the stick. Maintain aft stick until the
maximum pitch attitude is reached, which isindicated by the nose stopping and reversing
direction, and then push full forward on the stick to
generate a nosedown pitch rate. If the nosedown pitch
rate is high enough to break the deep stall, theaircraft will recover.
During some upright deep stalls, the aircraft may bestable with essentially no pitching motion. In thesecases, pull full aft stick (away from the ground) andmonitor nose movement. If nose movement occurs,continue stick cycling inphase. If nose movement is
not apparent after 34 seconds, then push full forward
on the stick to generate a nosedown pitch rate. This
nosedown pitch rate may be sufficient to reduce AOA
below the deep stall AOA and recover the aircraft. If
the nose does not continue down but reverses andstarts up, pull back on the stick and continue toreinforce these pitch cycles. Proper pitch rocking isaccomplished by allowing the nose to lead stickmotion; i.e., when nose movement reverses, the stickshould be reversed. When sufficient nosedown pitch
rate is generated to reduce the AOA below the deep
stall AOA, the aircraft will recover.
During upright deep stalls that are not stable, rolland yaw motions make it more difficult to determine
proper recovery inputs; however, pitch attitude is still
the best indication available. This pitch attitude isdetermined by the nose position with respect to thehorizon. If unable to determine pitch motions withoutside references, the ADI may be useful.
With proper stick cycling, the magnitude of the pitch
oscillations progressively increases until large
enough for recovery. Rapid fore and aft cycling of the
stick or cycling out of phase with the pitching motionof the aircraft will not be effective and may prevent
recovery. Pitch inputs must be abrupt and maximum
command. Pitch inputs that are smooth or less than
3129

T.O. GR1F16CJ1
maximum command do not generate pitch rate aseffectively, and may prevent recovery. Normally, only
one or two correctly applied cycles are required tobreak a deep stall; however, the presence of stores,
particularly a 300gallon fuel tank or 370gallon fueltanks, may necessitate five or more properly executed
stick cycles for recovery. Altitude loss is approximate
ly 10001500 feet per pitch rock cycle.

If still out of control:
Positive g, AOA indicator pegged at 32 degrees(upright deep stall) or negative g, AOA indicatorpegged at -5 degrees (inverted deep stall).
3.

If inverted, the same pitch rocking procedures apply
except if no pitch motion is apparent, the first stickcommand should be full forward (away from theground). Inverted deep stalls are generally stable,
regardless of the stores configuration. Yaw oscilla
tions may be noticed, but do not affect recovery.
If the pitch rate is still high as the aircraft recovers,
there may be a tendency for the aircraft to continue
pitching through to a deep stall in the other direction.
Attempt to stop the nose in a near vertical dive bytracking a spot on the ground. If the aircraft doestransition to an opposite AOA deep stall , it may bevery disorienting; however, pitch oscillations are
generally high and recovery should be rapid with afew properly executed stick cycles. Recovery is
confirmed by the nose remaining down and the AOA
remaining in the normal range. As the airspeed
increases above 200 knots, release the MPO switch,
maintain neutral roll and yaw commands, and apply
pitch commands as required to recover from theresulting dive using MIL/AB thrust.

MPO switch - OVRD and hold.
Maintain firm pressure.

FThe MPO switch must be held in the

OVRD position until the deep stall ispositively broken as evidenced by thepitch rate stopping, AOA in the normal
range (-5 to +25 degrees), and airspeed
increasing above 200 knots. Early
release of the MPO switch may delay
recovery.

FFailure to adequately secure andtighten lapbelt may result in inability
to reach and operate the MPO switch
during outofcontrol situations.

4.

Stick - Cycle inphase.

Pitch rocking with a high sustained
yaw rate may prevent recovery. Delaystick inputs until yaw rotation stops oris minimized. Pitch, roll, and yaw
oscillations associated with a deep
stall should not be confused with thecontinuous yaw rotation associatedwith a spin.

FRecovery from a deep stall condition

will present a low airspeed situation inwhich the aircraft may require morethan 6000 feet of altitude to attainlevel flight.

FIf recovery (pitch rate stopped, AOA

within -5 to +25 degrees, and airspeed
200 knots or greater) is not apparent
by 6000 feet AGL, eject.

The engine may stall when out of control. Also, FLCS
failure indications may occur. Ignore these indica
tions and concentrate on recovery.
In the event of a departure from controlled flight,
accomplish as much of the following as required toeffect a recovery:
1.

Controls - Release.

2.

Throttle- GE 129 IDLE, PW 229 MIL if in AB.
PW 229 If other than AB, do not move thethrottle.

3130OXYGEN MALFUNCTION PX IIThe OXY LOW caution light indicates oxygen
quantity below 0.5 liter or pressure below 42 psi.
If OXY LOW caution light illuminates:
1.

Cockpit pressure altitude - 10,000 feet maxi
mum.

If unable to descend immediately:
2.

Emergency oxygen - Activate.

3.

Oxygen hose - Disconnect.

T.O. GR1F16CJ1OBOGS MALFUNCTION PX III
An OBOGS malfunction may be indicated by either
an OXY LOW warning light or difficulty in breathing
through the oxygen mask. An OXY LOW warning
light is activated by low partial pressure of oxygen(PPO2), regulator pressure falling below 5 psi, or amonitor failure. Difficulty in breathing through themask may indicate an OBOGS malfunction combinedwith a failure in the warning light circuit. If difficultyin breathing through the mask occurs, activate EOS
if above 10,000 feet cockpit altitude, descend, andland as soon as practical.

6.

Altitude - Descend to cockpit altitude below
10,000 feet.

7.

Land as soon as practical.

If OXY LOW warning light begins flashing when BIT
is selected:
OBOGS monitor has failed.
5.

OBOGS BIT switch-BIT.
Returns OXY LOW warning light to steady.

6.

Altitude - Descend to cockpit altitude below
10,000 feet.

If OXY LOW warning light illuminates:

7.

Land as soon as practical.

1.

PBG MALFUNCTIONOXYGEN regulator pressure and cockpit
altitude-Check.

If pressure is less than 5 psi and cockpit altitude isabove 10,000 feet, or if pressure is greater than 5 psi
and cockpit altitude is above 25,000 feet:2.

EOS - Activate.

3.

Altitude - Descend to cockpit altitude below
10,000 feet.

4.

Land as soon as practical.

If pressure is less than 5 psi and cockpit altitude isbelow 10,000 feet:2.

Land as soon as practical.
Do not exceed cockpit altitude of 10,000 feet.

If pressure is greater than 5 psi and cockpit altitude
is below 25,000 feet:2.

Diluter lever - 100%.

If OXY LOW warning light goes off within 10 seconds:
Partial pressure of oxygen is sufficient for
operation in 100% but is not sufficient for operationin NORM.
3.

Continue mission with diluter lever in 100%.

If OXY LOW warning light remains on or diluter lever
was in 100% when light illuminated:
4.

OBOGS BIT switch-BIT.

If OXY LOW warning light remains on steady:Partial pressure of oxygen is not sufficient.
5.

EOS-Activate if cockpit altitude is above
10,000 feet.

A malfunction of the oxygen regulator while in PBG
may cause excessive pressure or failure of pressure to
decrease when g is reduced.
If excessive pressure is experienced or high pressure
continues after g is reduced:
1.

OXYGEN mode lever-ON.

If pressure is not relieved:2.

Oxygen hose-Disconnect.

3.

Cockpit pressure altitude-10,000 feet maxi
mum.

If unable to descend immediately:
4.

Emergency oxygen-Activate.

5.

Land as soon as practical.

SMOKE OR FUMESAll unidentified odors will be considered toxic. Do nottake off when unidentified odors are present. Do notconfuse ECS condensation for smoke.
If smoke or fumes are detected:
1.

OXYGEN REGULATOR - Check ON, 100%,and EMER.
NOTE
The emergency oxygen bottle is notrecommended for use in the smoke andfumes environment unless aircraft
oxygen supply contamination is sus
pected. Activation of the emergency
oxygen bottle does not prevent cockpit
smoke or fumes from entering theoxygen mask.
3131

T.O. GR1F16CJ1
2.

Altitude - 25,000 feet maximum.

3.

Airspeed - 500 knots maximum.

4.

AIR SOURCE knob-RAM.
External fuel cannot be transferred in OFF or
RAM. Consider jettisoning tank(s) todecrease drag if range is critical and the ECS
cannot be turned on for short periods of timeto transfer fuel.

PX III If AIR SOURCE knob is placed to

OFF or RAM, OBOGS is inoperative.
Activate EOS if OXY LOW warning
light illuminates above 10,000 feetcockpit altitude.
5.

Nonessential electrical equipment-Off.
NOTE
If in VMC and the ADI and HSI are notrequired for flight, the EGI/INS should
be considered nonessential.

6.

Determine cause of smoke or fumes andcorrect (if possible).

LANDING EMERGENCIES
Generally, the type of pattern flown in an emergency
is either an SFO or straightin approach and depends
on several factors:F

Nature of the emergency.

F

Weather conditions.

FDay or night.

FProximity of a suitable landing runway.

F

Fuel status.

STRAIGHTIN LANDING
A straightin landing is recommended for emergencies which dictate minimum maneuvering inputs
such as hydraulic, flight control, or electricalproblems or situations which result in a relatively
high thrust level being maintained to touchdown
such as a stuck or closed nozzle or when the engine isoperating satisfactorily in SEC. A controllability
check should be accomplished prior to commencing
the approach if minimum flying airspeeds or control
difficulties are experienced or are anticipated.

NOTE
FSmoke in the cockpit may be indicativeof an engine oil system malfunction. If
possible, retard throttle to lowest
setting possible to sustain flight andmonitor the OIL pressure indicator.
Refer to OIL SYSTEM MALFUNC
TION, this section, if appropriate.

FAny odor that smells of burning flesh

may be indicative of bird ingestion into
the engine. Monitor engine instru
ments for signs of abnormal operation.

7.

Land as soon as possible.

If cockpit visibility precludes safe operation:
8.

Airspeed-180 knots maximum.

9.

Seat-Full down.

10. ALT FLAPS switch-EXTEND.
11.

Canopy-Jettison.

3132

NOTE
When landing in SEC, an increased
ground roll distance is required due tohigher idle thrust.
SIMULATED FLAMEOUT (SFO) LANDING
Anytime engine failure is anticipated (abnormal
engine response, oil system failures, low fuel, etc.), an
SFO landing should be performed. At or just prior tohigh key, turn the EPU on, and if engine seizure is not
anticipated, turn the JFS on and verify theiroperation (EPU run and JFS RUN lights on). If theengine is still running at touchdown, the JFS shuts
down at WOW.
Fly the SFO pattern and landing in accordance withthe procedures for FLAMEOUT LANDING, thissection. If the engine fails, this action provides
sufficient energy to safely land the aircraft or to zoom
and eject if a safe landing cannot be made.

T.O. GR1F16CJ1An SFO landing is not recommended when landing
with the engine operating satisfactorily in SEC. The
higher level of idle thrust may result in a long andfast landing and difficulty stopping the aircraft.
To simulate an engine out glide with the LG up, use
idle thrust and 30 degrees speedbrakes. From thefront cockpit, this equates to the intersection of the
top of the speedbrakes and a line drawn from the tipof the horizontal tail to the top of the vertical tail root
fairing. To simulate an engine out glide with the LG
down, use idle thrust and 20 degrees speedbrakes.
From the front cockpit, this equates to theintersection of the top of speedbrakes and a line
drawn from the tip of the horizontal tail to the base of
the vertical tail root fairing. The additional dragproduced during an engineout condition is equiva
lent to retaining stores with a drag index of 170 with
the LG up or 70 with the LG down. If stores are
retained, adjust speedbrake deflection accordingly. If
the engine fails, close speedbrakes and jettison
stores.
When flying an SFO approach with the engine
operating at a higher thrust level than normal idle,
control descent rate and airspeed with the speedbra
kes rather than adjust the ground track. If thrust isexcessively high or if landing on a runway where
stopping distance may be critical, the procedures for
ABNORMAL ENGINE RESPONSE, this section,
should be considered.

On runways with less than 8000 feet
and without arresting gear or dragchute, there may be insufficientdistance to safely stop the aircraft.
After touchdown from an SFO landing, use a normal or
short field stopping technique as required by the
stopping distance available. Extend the hook if
required. If the engine rpm is greater than normal withthe throttle in IDLE or some other malfunction requires
excessive braking action to maintain a safe taxi speed,
the brakes may absorb a high amount of energy in ashort period of time. If required, refer to HOT BRAKES,
this section.

The following items should be accomplished:1.

Attain safe altitude.
NOTE
In the event that structural damage of
unknown extent is encountered or if
continued control of the aircraft is in
doubt, consider accomplishing applicablesteps of EJECTION (TIME PERMIT
TING), this section, prior to proceedingwith CONTROLLABILITY CHECK.

2.

GW-Reduce (as required).

3.

LE FLAPS switch-LOCK (if required).
If LEF damage is observed, consider locking
LEF's.

4.

Determine optimum configuration available
for landing.

If a condition which might cause
asymmetric TEF extension exists, consid
er alternate LG extension with the LG
handle in UP to preclude TEF extension.
If the LG handle remains up:
FFinal approach airspeed is 20
knots higher than normal.

F

The TO/LDG CONFIG warning
light may illuminate.

FPW 229 Nozzle remains closed,
resulting in higher than normal
landing thrust.

F

NWS is inoperative.

F

BRAKES
selected.

FFLCS remains in cruise gains.
Consider positioning AIR RE
FUEL switch to OPEN to obtaintakeoff and landing gains.

FThe LG handle warning lightremains on to indicate the position of
the gear handle is not in agreementwith the actual gear position.

CONTROLLABILITY CHECKWhen structural damage or any other failure that
may adversely affect aircraft handling characteris
tics is known or suspected, a controllability checkshould be performed.

CHAN

2

must

be

3133

T.O. GR1F16CJ1
5.

Stores-Selectively jettison (if required).
Refer to SELECTIVE JETTISON, this section.

6.

Slow only to that AOA/airspeed which allows
acceptable handling qualities.

FThe hook may miss the cable if theaircraft is not slow enough to compress
the MLG struts sufficiently to makeWOW or if forward stick pressure is held.

FTo prevent hook bounce and possible

missed engagement, avoid runway
centerline lighting.
NOTE

If the aircraft is not controllable down
to a reasonable landing speed (given
consideration to weather, runway
condition, facilities, pilot experience,
pilot arm fatigue, etc.), an ejection isrecommended.

FUnder certain conditions, arrestment

may produce a bouncing motion whichis readily apparent.

FOffcenter engagement results in aircraft yaw motions during cable runout.

CABLE ARRESTMENT
Refer to figure 55 for hook engagement limits. If
there is any doubt about stopping on the remaining
runway, lower the hook. Engage the cable as close tocenter as possible, nosewheel on the runway withbrakes off and aircraft aligned with the runway.
Place the HOOK switch to DN at least 1500 feet
before reaching the desired arresting cable andreduce speed as much as possible; however, if
brakes and NWS are inoperative, use flaperons andrudder as required to maintain directional control.
As the aircraft slows to below 70 knots, directional
control is reduced and the aircraft drifts right.
For most approachend arrestments, touchdown
should be at least 500 feet in front of the cable toallow sufficient time to lower the nosewheel to therunway prior to engagement. For an approachend
arrestment with one MLG up or damaged as
described in LANDING WITH LG UNSAFE/UP,
this section, maintain landing attitude after
touchdown and prior to engagement. Immediately
after touchdown, retard throttle to IDLE.
After engagement, rollback should be controlled by
the throttle. For cable disengagement, place HOOK
switch to UP and use approximately 1015 percent
increase in rpm to allow a rollback disengagement.

FUp to 5 seconds (after activation) are

required to fully raise the BAK14
cable.

1.

GW-Reduce (as required).

2.

HOOK switch-DN.

S Approachend arrestment:Touch down atleast 500 feet in front of the cable.

S Departureend arrestment:HOOK switch to

DN at least 1500 feet before reaching thecable.

3.

SHOULDER HARNESS knob-LOCKED.

4.

Consider options available if a missed
engagement occurs.

Prior to cable engagement:
5.

Throttle-IDLE.

6.

NWS-Engage (if required).

7.

Engage cable as close to center as possible;
nosewheel on the runway (if required) andbrakes off.

FCable arrestment at speeds greaterthan emergency arrestment speed,with offcenter distances greater than
35 feet, or with the nosewheel in the air
could result in structural failure of the
NLG, hook, and/or hook backup
structure.

3134

Using forward stick pressure to keep
an abnormally fast aircraft on the
runway for cable engagement willprobably result in a missed engage
ment or failure of the nose tire/NLG.

T.O. GR1F16CJ1

Do not use brakes while the cable isstretched or while being pulled back
ward. This action can result in aircraft
tipping backward. Control rollbackwith the throttle.
NET ARRESTMENT
Refer to NET ARRESTMENT LIMITATIONS,
section 5. Engaging a net barrier requires minimalpilot action as there is no hook to lower and little rollback after the aircraft's forward motion stops.
1.

SHOULDER HARNESS knob - LOCKED.

2.

Brakes - Release prior to engagement.

3.

Throttle - Off prior to engagement.

4.

Engage net perpendicular, preferably in thecenter portion of the runway.

The canopy should be retained
throughout the engagement to provide
pilot protection. Barrier netting will
not prevent subsequent canopy open
ing/jettison.

Engage net perpendicular to preclude
aircraft rotating sideways during thearrestment. Avoid steering back toward the center of the runway justprior to engagement as this could
result in a nonperpendicular engage
ment. Nosewheel steering is not
required; however, if engaged, it may
be left engaged. The throttle should be
retarded to off prior to engagement to
reduce the possibility of foreign object
damage.
LANDING WITH A BLOWN TIRE
When landing with a blown MLG tire, the landing
gear may collapse during landing roll if portions of
the tire remain and cause a wheel imbalance
condition. To avoid possible directional control

problems associated with landing gear collapse, anapproachend arrestment is preferred over a normal
approach and landing. To reduce the possibility ofdamage, the lowest practical landing GW andairspeed should be attained. Jettison stores if
possible. Retain empty external fuel tanks. Stores/
suspension equipment at stations 3 and/or 7 may
cause external fuel tanks at stations 4 and 6 to moveinboard when jettisoned. If an approachend
arrestment is not available and a normal approachand landing is flown, leave the antiskid system on
to minimize the possibility of skidding on the goodtire. If the wheel with the blown tire does not turnafter landing, the antiskid system switches to the b2t
alternate braking mode, LESS b2t pulsating antiskid
mode. Use roll control to relieve pressure on the blown
tire and NWS to maintain directional control.

With a blown tire, avoid centerline
lights as they may cause wheel damage
and subsequent loss of directional
control.
Stop straight ahead and shut down the engine as soon
as firefighting equipment is available. Do not attempt
to taxi unless an emergency situation exists.
If the blown tire is on the NLG, directional control
may be a problem due to a reverse castering effect. An
approachend arrestment with the nosewheel off the
runway is recommended. Depending on GW andspeed at the time of engagement, type of cable
engaged, and height of nosewheel above the runway,
it is possible for NLG strut failure and/or inlet
structural damage to occur. To reduce the possibilityof such damage, the lowest practical landing GW andairspeed should be attained. Jettison stores if
possible. Retain empty external fuel tanks. Stores/
suspension equipment at stations 3 and/or 7 may
cause external fuel tanks at stations 4 and 6 to moveinboard when jettisoned. Landing should be madewith the remaining internal fuel in the aft system.
After touchdown from a 13 degree AOA approach,
pitch attitude should be reduced to approximately 5
degrees prior to cable engagement. The HUD gun
borecross can be used to determine pitch attitude.
During the arrestment, the aircraft is likely to turnslightly right before stopping.
If an approachend arrestment is not available, referto procedures for aborting with a blown nose tire in
BLOWN TIRE ON TAKEOFF, this section, or
consider an all LG up landing (refer to LANDING
WITH LG UNSAFE/UP, this section).

3135

T.O. GR1F16CJ1Landing With A Blown Main Gear Tire

Landing With A Blown Nose Gear Tire

Prior to landing:Prior to landing:

1.

Stores - Jettison. Refer to JETTISON, thissection.
Retain empty external fuel tanks.

2.

GW-Reduce (if practical).

3.

TANK INERTING switch-TANK INERTING even if Halon is not available.

4.

AIR REFUEL switch-OPEN, if external fueltank(s) is installed.

Failure to depressurize external fueltank(s) significantly increases the prob
ability of tank explosion and fire if theaircraft departs the runway.
NOTE
Delay placing the AIR REFUEL switch
to OPEN until all external tanks are
empty.
5.

6.

7.

HOOK switch-DN.
An approachend arrestment is recom
mended. Refer to CABLE ARRESTMENT,
this section.

2.

GW-Reduce (if practical).
Plan to land with approximately 1500 poundsof fuel on board.

3.

Fuel distribution-All fuel in aft tank system(if practical).
At 3000 pounds fuel remaining, place ENG
FEED knob to FWD. When forward reservoir
is empty, place ENG FEED knob to NORM.
(Emptying forward tank system takes
approximately C 15 minutes, D 9 minutes if
fuel flow is 4000 pph. When forward tanksystem empties, the fuel in aft tank system isapproximately C 2000 pounds, D 2400
pounds.)4.

TANK INERTING switch - TANK INERTING
even if Halon is not available.

5.

AIR REFUEL switch-OPEN, if external fueltank(s) is installed.

Failure to depressurize external fueltank(s) significantly increases theprobability of tank explosion and fire if
the nose gear collapses during thearrestment.
6.

HOOK switch - DN.
An approachend cable arrestment with thenosewheel off the runway is recommended.
Refer to CABLE ARRESTMENT, this sec
tion.

7.

Final approach AOA - 13 degrees.

Final approach AOA-13 degrees.

NOTE

9.

Stores-Jettison. Refer to JETTISON, thissection.
Retain empty external fuel tanks.

ANTISKID switch-ANTISKID.
Use of antiskid minimizes skidding on goodtire during braking.

If a missed approachend cable arrestment occurs or
no approachend cable is available:

8.

1.

After touchdown:

If no approachend cable is available,
land on the side of runway away fromthe blown tire.

8.

NWS-Engage (if required).
The NWS light does not illuminate when
NWS is engaged if the AIR REFUEL switch
is in OPEN.

9.

Brake-As desired on good tire.

10. Maintain pitch attitude and go around.

3136

Stick - Lower nose to approximately 5 degrees
pitch attitude for arrestment.

After cable engagement:
Stick - Apply aft stick after nose starts down
to reduce load on the NLG.

If a missed cable engagement occurs:

T.O. GR1F16CJ1S Speedbrakes are not limited to 43 degrees.
With a blown NLG tire and loss of
NWS, it may not be possible to prevent
departure from the runway. A reverse
castering effect may occur in which thenosewheel moves opposite to therudder or differential braking input.
NOTE
The maximum allowable fuel flow with
one reservoir empty is 25,000 pph.
LG EXTENSION MALFUNCTIONS
Malfunctions in extending the LG are normallyindicated by failure of the LG handle to lower or by
failure of one or more LG's to extend accompanied bylack of the corresponding WHEELS down lights andcontinuous illumination of the LG handle warning
light. Refer to LG FAILS TO EXTEND, this section.
Mechanical failures can have varying causes. Failureof an LG component usually affects only one LG andcannot be corrected in flight. Abnormally highmoisture content in the hydraulic fluid can cause
more than one LG to fail to extend after prolonged
operation at low ambient temperatures. Thissituation can usually be corrected after several cyclesof the LG handle at low altitude.
Abnormal indications after the LG handle is lowered
can also be caused by electrically related malfunc
tions (i.e., electrical shorts, electrical component
failures, or cannon plug problems). The malfunction
may result in an indication problem or an actual
failure of one or more LG to extend. A visualconfirmation of the LG position should be obtained if
possible before any action is taken.
A common failure mode is electrical shorting withina downlock switch caused by moisture intrusion. This
shorting typically occurs after the LG handle islowered and results in opening of the LG uplock/downlock circuit breaker. Opening of this circuit
breaker causes all WHEELS down lights to remain
off. The light in the handle may go off normally or it
may remain on if the shorting has also adversely
affected the warning light circuit. Another result of
this circuit breaker opening is that the LG will go into
hydraulic isolation immediately after the LG handle
is raised and the LG won't retract. Other effects
associated with an open LG uplock/downlock circuit
breaker are as follows:

S No AOA bracket in the HUD.

S Landing/taxi light is inoperative.
S NWS is inoperative.
An indication failure may be distinguished from an
actual extension malfunction by the WHEELS down
lights and the LG handle warning light. The
WHEELS down lights and the LG handle warning
light have separate power sources and circuitry.
Thus, if either circuit operates correctly, the LGshould be down and locked even though the othercircuit may indicate a problem. If possible, cycle the
LG to reattempt extension while watching the
WHEELS down lights and the LG handle warning
light.
Because of the number of possible malfunctions,specific procedures for every situation are notfeasible. If time and conditions permit, groundsupervisory and technical assistance should berequested. Evaluate available options prior to using
the ALT GEAR handle. Other options may include an
LG up landing, diverting to a more suitable landing
field, landing where the hazards of departing the
prepared surface are minimal, or using a cablearrestment.
LG Handle Will Not Lower
If the LG handle cannot be moved to the DN position
after depressing the LG handle down permission
button, the electrical circuitry or solenoid has
probably failed. C DF The DN LOCK REL button
mechanically permits the LG handle to be moved to
the DN position. If the LG handle cannot be moved to
the DN position due to a mechanical failure, CHAN 1
brakes are inoperative and TEF's must be extended
using the ALT FLAPS switch PW 229 and the nozzle
remains closed. With the nozzle closed, idle thrust isapproximately 400 pounds greater than normal.
If LG handle cannot be lowered normally:
1.

DN LOCK REL button-Depress and lower
LG handle.

If LG handle still cannot be lowered:2.

ALT FLAPS switch-EXTEND.

3.

BRAKES channel switch-CHAN 2.

4.

Go to ALTERNATE LG EXTENSION, thissection.
PW 229 Nozzle remains closed, resulting inhigher than normal landing thrust.
3137

T.O. GR1F16CJ1brace assembly may not be functioning
properly. The LG may appear down, but
the LG may collapse during landing.
Plan on using the LG unsafe/up
procedures even if the LG eventuallyindicates normal. Refer to LANDING
WITH LG UNSAFE/UP, this section.

NOTE
After a successful alternate gear
extension with the landing gear handle
still up, the LG handle warning lightremains on to indicate the position of
the gear handle is not in agreementwith the actual gear position.
LG Fails To Extend

If the LG previously failed to retract,
do not cycle the LG handle. Damage tothe LG or LG doors may preclude
successful extension.

NOTE
If alternate LG extension was per
formed and one or more LG indicate
unsafe, refer to ALTERNATE LG
EXTENSION, this section.

NOTE

If one or more LG fail to extend, the LG handle may
be cycled multiple times to attempt to extend the LG
if no structural/battle damage exists and if normal
hydraulic pressure exists. Multiple cycle attempts
may release the locking pawls and result in normal
gear extension if the failure to extend was caused by
failure of the uplocks to release.

If the NLG WHEELS down light is off,
confirmation of the NLG position can bemade by checking landing/taxi light
operation. Illumination of either light
confirms that the NLG is down. With the
NLG WHEELS down light off, NWS
may be inoperative (without a NWS
FAIL caution light).
1.

If the LG previously failed to retract,
do not cycle the LG handle. Damage tothe LG or LG doors may preclude
successful extension.
If cycling of the LG handle fails to correct thesituation, use of the alternate extension system must
be considered.

If LG handle warning light came on when the LG
handle was lowered, then went off, and tests good or
if WHEELS down lights operated normally:
2.

If possible, get a visual confirmation of LG position.
If the NLG WHEELS down light is off, confirmationof the NLG position can be made by checking
landing/taxi light operation. Illumination of either
light confirms that the NLG is down. With the NLG
WHEELS down light off, NWS may be inoperative(without a NWS FAIL caution light).

3.

3138

Speedbrakes - Verify opening is less than 43
degrees.
From the front cockpit, the top of the
speedbrakes should be slightly above a line
drawn from the tip of the horizontal tail to the
top of the vertical tail root fairing.

If RMLG WHEELS down light is off,
speedbrakes may not be limited to 43
degrees.

If one or more LG indicate unsafe:

If at anytime, an LG intermittently
indicates unsafe (i.e. WHEELS down
light off and LG handle warning lighton), the overcenter lock on the LG dragLG handle-Cycle and monitor LG handle
warning light and WHEELS down lights.

Land normally.

If LG handle warning light did not illuminate or
remained illuminated after LG handle was lowered and
if one or more WHEELS down lights did not illuminate:
4.

Go to ALTERNATE LG EXTENSION, thissection.

T.O. GR1F16CJ1
Alternate LG Extension

If LG indicates safe:

Alternate LG extension should be accomplished at
the lowest practical airspeed below 190 knots. The
NLG may not indicate down until airspeed is reduced
below 190 knots. NWS is inoperative after alternate
LG extension even if system B hydraulic pressure isavailable.

3.

Land normally.
If possible, get visual confirmation of LG
position. If all WHEELS down lights were
initially off with the LG handle down anduse of the hook may be required after
touchdown, verify before landing that thehook extends.

1.

4.

Stop straight ahead on the runway.

LG handle-DN. (Use DN LOCK REL, if
required.)FNWS is not available following alter

nate LG extension.

FDo not delay lowering LG below 2000

feet AGL.

FIf the LG was alternately extended due

to failure of system B, only brake/JFS
accumulator braking is available andafter stopping, the parking brakeshould be engaged until chocks are
installed.

FIf LG handle does not lower, selectBRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND. PW 229
Nozzle remains closed, resulting inhigher than normal landing thrust.

2.

ALT GEAR handle-Pull (if required). (190knots maximum, if practical).

If LG indicates unsafe:
3.

S Alternate LG extension can be used up to 300

knots; however, the NLG may not fully extend
until 190 knots. Time above 190 knots should
be minimized in case there is a leak in thepneumatic lines.

If LG indicates safe:
4.

Land normally.
If possible, get visual confirmation of LG
position.

5.

Stop straight ahead on the runway.

S If an unsafe MLG indication exists and both

MLG are out of the wheel wells, pulling the
ALT GEAR handle is not recommended.

If the LG was alternately extended due
to failure of system B, only brake/JFS
accumulator braking is available andafter stopping the parking brakeshould be engaged until chocks are
installed.

FNWS is not available following alter

nate LG extension.

FDo not depress the ALT GEAR reset

button while pulling the ALT GEAR
handle. This action may preclude
successful LG extension.

FPulling the ALT GEAR handle withnormal system B hydraulic pressure,e.g., NLG fails to extend, may result inhydraulic system B failure within 15
minutes.

Stick-Apply alternating g forces (-1.0 to +3.0g) to free LG.
Up to 300 knots may be required to provide
sufficient g force.

If LG still indicates unsafe:4.

Speedbrakes - Verify opening is less than 43
degrees.
From the front cockpit, the top of the
speedbrakes should be slightly above a line
drawn from the tip of the horizontal tail to the
top of the vertical tail root fairing.
3139

T.O. GR1F16CJ1
NOTE
If RMLG WHEELS down light is off,
speedbrakes may not be limited to 43
degrees.
5.

Go to LANDING WITH LG UNSAFE/UP, thissection.

Landing With LG Unsafe/Up
Prior to landing with any of the LG unsafe or up,
consider the following:

S Airfield facilities.
S Hook engagement limits.
S Crosswind component.
S Runway and overrun conditions.
If conditions are not favorable:
1.

Refer to EJECTION (TIME PERMITTING),
this section.

To accomplish the landing:1.

Retain empty fuel tanks and racks.

Delay placing the AIR REFUEL switch
to OPEN until all external fuel tanks
are empty.
6.
7.

FCR-OFF.
PX II SMS PX III ST STA/HDPT/ECM power -

Off.

8.

SHOULDER HARNESS knob-LOCKED.

9.

Refer to figure 316.
NOTE
If either MLG is not extended, EPU
operation cannot be terminated with
the EPU switch after engine shutdown.

ANTISKID MALFUNCTION (LANDING)
Illumination of the ANTI SKID caution light whenthe LG handle is lowered indicates one of thefollowing:

S b2t Loss of power to one of the two brake channels
or from the ANTISKID switch.

S b2t Built in Test (BIT) detected malfunction of oneof the two brake channels.

S LESS b2t Failure in the antiskid system (pulsating
If time permits, delay landing until
external fuel tanks are empty. If an
immediate landing is required, jetti
son all external fuel tanks.
2.

Armament-Jettison.

3.

GW-Reduce.

4.

TANK INERTING switch-TANK INERTING even if Halon is not available.

5.

AIR REFUEL switch-OPEN.

Failure to depressurize external fueltanks significantly increases the prob
ability of tank explosion and fire.

3140

brake pressure mode activated).

S LESS b2t Loss of electrical power to the CHAN 1
brake and antiskid control circuits.

b2t Cycling the ANTISKID switch to OFF and backto ANTISKID will not extinguish the ANTI SKID
caution light. Touchdown skid control protection
may not be available, so do not apply brakes before
touchdown.
b2t If the ANTI SKID light extinguishes on wheel

spin up, antiskid protection is available and one of
the two brake control channels is fully functional.
Normal braking may be used. If the light does not
extinguish on wheel spinup, or first illuminates on
wheel spinup, braking may be in the alternate
braking mode, detectable by a pulsing sensation
when differential braking is applied. In the alternate
braking mode, symmetric braking will give the best
stopping performance.

T.O. GR1F16CJ1LESS b2t Changing to CHAN 2 brakes may
extinguish the ANTI SKID caution light and provide
normal antiskid and braking functions. If the failure
was in the antiskid system, cycling the ANTISKID
switch may correct the problem. If the ANTI SKID
caution light remains on, antiskid protection is notavailable and degraded (i.e. pulsating) or inopera
tive braking should be expected.
If the ANTI SKID caution light illuminates (with the
ANTISKID switch in ANTISKID) when the LG
handle is lowered:
1.

BRAKES channel switch-CHAN 2.

b2t If the ANTI SKID caution light remains on:

2.

Refer
to
ANTISKID(GROUND), this section.

MALFUNCTIONLESS b2t If the ANTI SKID caution light remains on:
2.

ANTISKID switch-OFF, then back to ANTISKID.

If the ANTI SKID caution light remains on, expect
pulsating or inoperative braking and no antiskid
protection during landing:
3.

Refer
to
ANTISKID(GROUND), this section.

MALFUNCTION

ASYMMETRIC STORES (LANDING)
Refer to figure 317 for computation of asymmetricmoment. A modified approach is required if the net
asymmetry exceeds 10,000 footpounds. Successful
landings have been demonstrated with asymme
tries as large as 25,020 footpounds. If the net
asymmetry exceeds 25,020 footpounds, storesshould be selectively jettisoned from the heavy wing
to reduce the asymmetry. The decision to land withlarge asymmetries should consider such factors as
weather conditions, runway length/width, surface
conditions (RCR), arresting gear availability,
crosswind component/gusts, and pilot experience.
Avoid abrupt control inputs. Limit maximum bankangle changes to 90 degrees and do not exceed 10
degrees AOA until the net asymmetry can be
determined. A controllability check should be
performed with the LG down to determine handling
qualities and the feasibility of landing. During thecontrollability check, determine maximum maneu

vering AOA by slowly increasing AOA until rollauthority is insufficient to maintain wings level. Do
not exceed 12 degrees AOA during this check. If lossof roll authority is experienced, the maximum
maneuvering AOA is 2 degrees below that at whichroll authority was lost or 10 degrees AOA,
whichever is less. If landing is feasible, plan to flya shallow, poweron, straightin approach. Use rolltrim and lateral stick as required. Full roll trim may
not be enough to compensate for large asymmetries.
Trim rudder into the heavy wing up to a maximumof two dots. This decreases bank angle and permits
a low bank angle, low sideslip approach, minimiz
ing pilot induced lateral oscillations. This action
increases the roll trim necessary to hold up theheavy wing.
Fly a shallow, poweron, straightin approach andreduce bank angle to wings level on short final. Do
not exceed the maximum maneuvering AOA, as
determined during the controllability check, on
final approach or during the flare/touchdown.
Twopoint aerodynamic braking should be per
formed at 1011 degrees AOA until approximately
120 knots. The nose should then be lowered to therunway and maximum effort antiskid braking
should be used. At no time should 11 degrees AOA
be exceeded while on the runway. When the ARI
switches out with wheel spinup, large rudderchanges can occur which cause yaw into the heavy
wing. When landing with no crosswind, this yaw
helps align the aircraft with the runway. Withcrosswind components greater than 10 knots (5knots if the net asymmetry exceeds 20,000
footpounds), land with the heavy wing into thecrosswind even if this results in landing downwind.
Failure to do so may result in inadequate roll
control.
1.

Compare weights on mirror stations (4 and 6,
3 and 7, etc.). Include wingtip AIM120's;
exclude wingtip AIM9's since their weight isoffset by the lift they generate.

2.

Determine asymmetric moment for each set of
stations. Enter with weight difference atbottom of chart, proceed vertically to theappropriate line, and proceed horizontally left
to read the asymmetric moment.

3.

Add or subtract each asymmetric moment todetermine net asymmetry.

3141

T.O. GR1F16CJ1LG Unsafe/Up Landing
APPROACHEND ARRESTMENTCONFIGURATIONAVAILABLE

ALL LG INDICATE UNSAFE
BUT APPEAR NORMAL

10. HOOK-DOWN.
11. APPROACHEND CABLE-ENGAGE.

ALL LG UP

BOTH MLG UP OR UNSAFEUNAVAILABLE
10. LAND NORMALLY.

ARRESTMENT NOT RECOMMENDED.
USE APPROACHEND ARRESTMENTUNAVAILABLE PROCEDURE.

10.
11.
12.
13.

ALT GEAR HANDLE-IN.
WAIT 5 SECONDS.
LG HANDLE-UP.
ALT GEAR RESET BUTTON-DEPRESS(2 SECONDS).
14. USE ALL LG UP PROCEDURE.
15. IF NLG DOES NOT RETRACT:a. HOOK-DOWN.
b. LOW ANGLE APPROACH AT 11°
AOA.
c. ATTEMPT A FLYIN ENGAGEMENT.
d. THROTTLE-OFF AFTER ENGAGEMENT.

10.
11.
12.
13.

10.
11.
12.
13.

EPU-ON.
ALT FLAPS-EXTEND.
LOW ANGLE APPROACH AT 13° AOA.
THROTTLE-OFF IMMEDIATELYPRIOR TO TOUCHDOWN.

ALT GEAR HANDLE-IN.
WAIT 5 SECONDS.
LG HANDLE-UP.
ALT GEAR RESET BUTTON-DEPRESS(2 SECONDS).
14. USE ALL LG UP PROCEDURE.
15. IF NLG DOES NOT RETRACT:
a. CONSIDER LANDING FROM A
LOW ANGLE APPROACH AT 13°
AOA IF WING FUEL TANKS ARECARRIED.
b. RECOMMEND EJECTION IF WING
FUEL TANKS ARE NOT CARRIEDOR IF CONDITIONS ARE NOTCONSIDERED FAVORABLE FOR
AN ATTEMPTED LANDING WITH
WING FUEL TANKS.

IF THE ENGAGEMENT ISMISSED, MAINTAIN WINGS LE
VEL AND GO AROUND. IF A
GOAROUND IS NOT ACCOMPLISHED, THE AIRCRAFT
MAY GROUND LOOP.
NLG UP OR UNSAFE

ARRESTMENT NOT RECOMMENDED.
USE APPROACHEND ARRESTMENTUNAVAILABLE PROCEDURE.

Figure 316.(Sheet 1)

3142

10. EPU-ON.
11. LOW ANGLE APPROACH AT 13° AOA.
12. THROTTLE-OFF AFTER TOUCH
DOWN.
13. LOWER NOSE TO RUNWAY BEFORECONTROL EFFECTIVENESS BEGINSTO DECAY.
14. EPU-OFF AFTER STOP.

T.O. GR1F16CJ1LG Unsafe/Up Landing
APPROACHEND ARRESTMENTCONFIGURATIONAVAILABLE

ONE MLG AND NLG UP OR UNSAFE

ONE MLG INDICATES UNSAFE
BUT APPEARS NORMALARRESTMENT NOT RECOMMENDED.
USE APPROACHEND ARRESTMENTUNAVAILABLE PROCEDURE.

10. HOOK-DOWN.
11. LOW ANGLE APPROACH AT 11° AOA.
12. AFTER TOUCHDOWN, USE ROLLCONTROL, IF NECESSARY, TO HOLD
WING UP. IF ROLL CONTROL IS NEEDED TO HOLD WING UP, MAINTAINLANDING ATTITUDE FOR ENGAGE
MENT. IF ROLL CONTROL IS NOTNEEDED TO HOLD WING UP, LOWERNOSE FOR ARRESTMENT.
13. THROTTLE-OFF AFTER ENGAGEMENT.

UNAVAILABLE
10.
11.
12.
13.

ALT GEAR HANDLE-IN.
WAIT 5 SECONDS.
LG HANDLE-UP.
ALT GEAR RESET BUTTON-DEPRESS(2 SECONDS).
14. USE ALL LG UP PROCEDURE.
15. IF LG DOES NOT RETRACT:
a. CONSIDER LANDING FROM A
LOW ANGLE APPROACH AT 13°
AOA IF EXTERNAL FUEL TANK(S) ISCARRIED.

NOTELAND ON THE SIDE OF THE
RUNWAY AWAY FROM THE UN
SAFE MLG.
b. RECOMMENDEJECTION
IF
EXTERNAL FUEL TANK(S) IS NOTCARRIED OR IF CONDITIONS ARENOT CONSIDERED FAVORABLE
FOR AN ATTEMPTED LANDING
WITH EXTERNAL FUEL TANK(S).

IF THE ENGAGEMENT IS MISSEDAND ROLL CONTROL WAS NECESSARY TO HOLD WING UP,
MAINTAIN WINGS LEVEL AND
GO AROUND. IF A GOAROUNDIS NOT ACCOMPLISHED, THE
AIRCRAFT MAY GROUND LOOP.
ONE MLG UP

10.
11.
12.
13.

ALT GEAR HANDLE-IN.
WAIT 5 SECONDS.
LG HANDLE-UP.
ALT GEAR RESET BUTTON-DEPRESS(2 SECONDS).
14. USE ALL LG UP PROCEDURE.
15. IF LG DOES NOT RETRACT:a. HOOK-DOWN.
b. LOW ANGLE APPROACH AT 11°
AOA.
c. AFTER TOUCHDOWN, USE ROLLCONTROL TO HOLD WING UP
AND MAINTAIN LANDING ATTITUDE FOR ENGAGEMENT.
d. THROTTLE-OFF AFTER ENGAGEMENT.

IF THE ENGAGEMENT IS MISSED,MAINTAIN WINGS LEVEL AND
GO AROUND. IF A GOAROUNDIS NOT ACCOMPLISHED, THE
AIRCRAFT MAY GROUND LOOP.

Figure 316.(Sheet 2)

3143

T.O. GR1F16CJ1

Asymmetric Moment

ASYMMETRIC MOMENT1000 FOOT-POUNDS

40

30

20
D
B
10

0

0

1

C

WEIGHT

2

A

3

4

1000 POUNDSGR1F-16CJ-1-0128X37

Figure 317.

S

SAMPLE PROBLEM 1:

S

Full 370gallon fuel tank on station 4; empty
fuel tank on station 6.

S

Three MK 82 (SNAKEYE) bombs on station 3,
station 7 empty.

A.

Station 4/6 weight
differenceB.

Asymmetric moment
footpounds

C.

Station 3/7 weight
differenceD.

Asymmetric moment= 16,500 footpoundsE.

Net asymmetry (B+D)= 30,500 footpounds= 2405 pounds= 14,000

Follow same procedures (A through D) as above;
however, since the asymmetric moments are on the
opposite wing, the moments are subtracted rather
than added.
E.

Net asymmetry (B-D) = 2500 footpounds

1.

AOA-10 degrees maximum.

= 1650 pounds
Large asymmetric loads severely limit
lateral control when rolling away fromthe heavy wing. Until determining net
asymmetry, limit maximum bankangle change to 90 degrees, avoid
abrupt control inputs, and do not
exceed 10 degrees AOA.

SAMPLE PROBLEM 2:

S

Full 370gallon fuel tank on station 6; empty
fuel tank on station 4.

3144

Three MK 82 (SNAKEYE) bombs on station 3;
station 7 empty.

2.

Determine net asymmetry.
Refer to figure 317.

T.O. GR1F16CJ1
If asymmetry is greater than 25,020 footpounds:
3.

Stores-Jettison (as required).
Selectively jettison stores from the heavy
wing to obtain a net asymmetry less than
25,020 footpounds. Refer to SELECTIVEJETTISON, this section.

7.

Rudder trim-Trim into the heavy wing (ifrequired).

If landing is not feasible:
5.

Go to EJECTION (TIME PERMITTING), thissection.

If asymmetry is greater than 10,000 footpounds:

If asymmetry is less than 10,000 footpounds:

4.

Controllability-Check.

4.

S Lower LG at a safe altitude and check

NLG WOW SWITCH FAILURE

handling qualities until roll authority is
insufficient or up to 12 degrees AOA
maximum.

S Maximum maneuvering AOA for approachand landing is 10 degrees AOA or 2 degrees
less than the AOA at which roll authority is
insufficient to maintain wings level, which
ever is less.

If landing is feasible:

The decision to land with a large
asymmetry should consider such factors as weather conditions, runway
length/width and surface conditions(RCR), arresting gear availability,
crosswind component/gusts, and pilot
experience.
5.

Land normally.

If the NLG WOW switch fails to the ground position,
the A/R DISC function on the stick is inoperative, thespeedbrakes are not limited to 43 degrees with theright MLG down, NWS can be engaged in flight, and
AOA indications are approximately 0 degrees.
1.

NWS-Engage.

If AR/NWS light comes on:
2.

NWS-Disengage.

3.

AR/NWS light-Off.
NOTE
Insure that AR/NWS light is off prior tolanding so that the NWS does notfollow rudder commands when thenosewheel is lowered to the runway.

4.

Fly a shallow, poweron, straightin approach.

Speedbrakes-Close to less than 43 degrees.
From the front cockpit, the top of the
speedbrakes should be slightly above a line
drawn from the tip of the horizontal tail to the
top of the vertical tail root fairing.

FWith crosswind component greaterthan 10 knots (5 knots if the net
asymmetry exceeds 20,000 footpounds), land with heavy wing into thecrosswind even if this results inlanding downwind. Failure to do so
may result in inadequate roll control.

FDo not exceed the maximum AOA, as

determined during the controllability
check, during final approach, flare,
touchdown, or twopoint aerodynamic
braking.

6.

Roll trim and lateral stick-As required.

Visually confirm speedbrake opening
is limited to 43 degrees to prevent thelower surfaces from striking the
runway during landing.
DRAG CHUTE FAILURE
If decision is made to goaround:1.

Drag chute-Release.

2.

Throttle-MAX AB.

3145/(3146 blank)

T.O. GR1F16CJ1

SECTION IV
CREW DUTIES(Not Applicable)

41/(42 blank)

T.O. GR1F16CJ1

SECTION V
OPERATING LIMITATIONSTABLE OF CONTENTS
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Instrument Markings . . . . . . . . . . . . . . . . . . . . . .
Engine Limitations . . . . . . . . . . . . . . . . . . . . . . . .
Fuels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
System Restrictions . . . . . . . . . . . . . . . . . . . . . . . .
Terrain Following Radar . . . . . . . . . . . . . . . . .
Jet Fuel Starter Limits . . . . . . . . . . . . . . . . . .
Tire Speed Limit . . . . . . . . . . . . . . . . . . . . . . . .
Brake Energy Limits . . . . . . . . . . . . . . . . . . . .
Fuel System Limitations . . . . . . . . . . . . . . . . .
One Reservoir Empty . . . . . . . . . . . . . . . . .
Negative G Flight . . . . . . . . . . . . . . . . . . . . .
Airspeed Limitations . . . . . . . . . . . . . . . . . . . . . .
System Airspeed Limitations . . . . . . . . . . . . .
Maximum Airspeed Operating
Limitations . . . . . . . . . . . . . . . . . . . . . . . . . .
Low Airspeed Operating Limitations . . . . . .
Prohibited Maneuvers . . . . . . . . . . . . . . . . . . . . .
Gross Weight Limitations . . . . . . . . . . . . . . . . . .
CenterofGravity Limitations . . . . . . . . . . . . . .
Acceleration Limitations . . . . . . . . . . . . . . . . . . .
AOA and Rolling Limitations . . . . . . . . . . . . . . .
Stores Limitations . . . . . . . . . . . . . . . . . . . . . . . . .
Asymmetric Stores Loading . . . . . . . . . . . . . .
Miscellaneous Limitations . . . . . . . . . . . . . . . . . .
Net Arrestment Limitations . . . . . . . . . . . . . .
Crosswind Limits . . . . . . . . . . . . . . . . . . . . . . .

Page
51
51
51
51
52
52
52
52
52
52
52
52
52
52
52
53
53
53
53
53
54
54
54
55
55
55

INTRODUCTION
The aircraft and system limitations that must be
observed during normal operations are presented
in this section and T.O. GR1F16CJ12. Thoselimitations that are characteristic of a special phaseof operations such as emergency procedures, etc.,
are not covered in this section.

INSTRUMENT MARKINGS
Refer to figure 51 for location and range of themarkings.

ENGINE LIMITATIONS
Refer to figure 52 for Engine Limitations and figure
53 for Engine-Operational Envelope.

FUELS
JP4, JP5, JP8; NATO F34, F35, F40, F43, andF44; and commercial JET A, JET A1, and JET B
are approved fuels. Except for freeze point andpossible icing and corrosion inhibitor differences,JP4, NATO F40, and JET B are equivalent and thesame operating limitations apply.
Operating and throttle movement limitations forfuels other than JP4, NATO F40, and JET B, arethe same as for JP4, NATO F40, and JET B
except:Ground starts with temperature below-4_F (-20_C) may produce more smoke and requirea longer time for engine lightoff. Ground starts
should not be attempted with fuel temperature
below -40_F (-40_C) except JET A which is -35_F(-37_C). Airstart lightoff times also may beslightly longer.
NOTE
Changing operations fromJP4, NATO F40, or JET B may resultin an increase in the number of AB
recycles.
PW 229PW 229 Fuels with very high flash points (JP5,
NATO F43, and NATO F44) may leave visible
signature on AB cancellation at high altitude.

Approved fuels may be intermixed in any
proportion during ground or AR operations. Nochange in engine operating limitations is required.
Due to fuel freeze points, non JP4 or NATO F40
fuel in external tanks may not transfer after
sustained operation (5 minutes or longer) below 275
knots from 25,00030,000 feet or below 0.72 machfrom 30,00042,000 feet (all non JP4 or NATO F40
fuels except JET A) or below 300 knots from25,00030,000 feet or below 0.83 mach from30,00042,000 feet (JET A).
NATO F34, NATO F40, and NATO F44 may notcontain corrosion inhibitor and NATO F35, NATOF43, JETA, JET A1, and JET B may not containicing or corrosion inhibitors. Restrict operation
without icing inhibitor to one flight. Restrict engine
operation without corrosion inhibitor to 10
consecutive hours.
51

T.O. GR1F16CJ1
Certain fuels are heavier; refer to FUEL QUANTITY
INDICATION AND TANK ARRANGEMENT, Section I.

SYSTEM RESTRICTIONSTERRAIN FOLLOWING RADAR
Refer to TF DISPLAYS AND OPERATION, T.O.
GR1F16CJ3411.

FUEL SYSTEM LIMITATIONSOne Reservoir Empty
The maximum allowable fuel flow with one reservoir
empty is 25,000 pph.
Negative G Flight

JET FUEL STARTER LIMITS

Negative g flight with both reservoir tanks full islimited to:Refer to figure 54.

SAB thrust - 10 seconds.

TIRE SPEED LIMITSMIL thrust or below - 30 seconds.

The MLG tires are certified for use to 225 knots
groundspeed. The NLG tire is certified for use to PX II
217, PX III 235 knots groundspeed.
BRAKE ENERGY LIMITS
Refer to T.O. GR1F16CJ11, Part 2 for brakeenergy limits for maximum effort braking, taxi,
aborted takeoff, landing, and the effect onturnaround capability. The actual energy per brake
may differ considerably from the value found in Part
2. This is caused by unequal energy distribution
between the brakes or residual heat from previous
braking. Maximum brake application speed is the
maximum speed from which the aircraft can be
stopped using maximum braking. This speed isbased on the capability of each brake to absorb amaximum of 23.5 million footpounds of energy.
(Refer to T.O. GR1F16CJ11, Part 2 ABORTEDTAKEOFF MAXIMUM BRAKE APPLICATIONSPEED.)NOTE
Negative g flight should be avoided
when a low fuel condition exists(forward or aft reservoir not full) or
ENG FEED knob out of NORM.

AIRSPEED LIMITATIONS

Maximum airspeed operating limita
tions may be easily exceeded in level
flight due to the improved performanceengine excess thrust capabilities.
SYSTEM AIRSPEED LIMITATIONSFInitiating maximum effort braking

above maximum brake application
speed may result in loss of braking
before the aircraft is stopped.

FDanger zone procedures should be

followed for any event which requires
excessive braking.

FIf brake energy absorption is in thedanger zone, wheel fusible plugs
release tire pressure within 315
minutes after the stop.

52Change 1

Refer to figures 53, 55, and 56.
MAXIMUM AIRSPEED OPERATING LIMITATIONS
Refer to figure 53. Maximum operating airspeed is
800 knots from sea level to 30,000 feet MSL. Above
30,000 feet MSL, the aircraft is limited to 2.05 mach.
Refer to figure 56 and T.O. GR1F16CJ12.
Maximum operating airspeed/mach may be reduced
as a result of system restrictions or stores limitations.

T.O. GR1F16CJ1LOW AIRSPEED OPERATING LIMITATIONS

3.

Recovery should be initiated no later than activationof the low speed warning tone.

Consecutive 360degree maximum commandrolls.

4.

Maximum command rolling maneuvers above
1.8 mach and either above 3g or below 35,000
feet MSL.

5.

Rudder rolls or rudderassisted rolls of morethan 90 degrees of bank angle change with any
store on station 3, 4, 6, or 7.

6.

With LG and/or TEF's down:

For CAT I configurations with dragindices greater than 120, delaying
recovery until activation of the low
speed warning tone may result in
departure regardless of recovery tech
nique. Rapid airspeed decay may
reduce control authority to the pointthat recovery inputs have no effect.
Low thrust settings, external fueltanks, or inlet pods increase thepossibility of a departure.
To avoid departures due to roll coupling, do not
operate with category III loadings below 200 knots
except for takeoff and landing.

Departures from controlled flight withasymmetric category III loadings may
result in fast, flat (possibly nonrecov
erable) spins.

PROHIBITED MANEUVERSThe following maneuvers are prohibited:1.

station 3, 4, 6, or 7.

S Maximum command rolls of more than 90
degrees of bank angle change.

7.

Abrupt roll reversals above 550 knots andbelow 20,000 feet MSL with the STORESCONFIG switch in CAT I.

8.

PX III With CFT's installed, abrupt rollreversals above 500 knots with the STORESCONFIG switch in CAT I.

9.

LESS b6n Rapid rudder release or reversal
above 300 knots/0.6 mach.

10. LESS b6n PX III Flight with CFT's installed andCAT III configurations.

GROSS WEIGHT LIMITATIONS
The maximum allowable GW for ground handling,taxi, takeoff, inflight, and landing is 48,000 pounds.

Intentional departures and spins with any of
the following:

CENTEROFGRAVITY LIMITATIONSS Symmetric category I loading with suspension

Refer to T.O. GR1F16CJ52.

S Asymmetric category I loading.

Generally, the aircraft is within CG limits when the
red portion of the AL pointer is not visible (FUELQTY SEL knob in NORM).

equipment or missiles at station 3, 4, 6, or 7.

S Category III loading.
S Altitude below 30,000 feet AGL.
S CG aft of aft limit for the configuration being
flown.

S Lateral fuel (internal and external) imbalance
greater than 200 pounds.

2.

S Flight above 15 degrees AOA with stores at

Repeated maximum rudder reversals.

ACCELERATION LIMITATIONS
Acceleration limitations are contained in figure 57this section, and T.O. GR1F16CJ12, figures 59
and 510.
Load factor limits should not be intentionally
exceeded. Notify maintenance of a possible overg if
symmetric maneuvering on the g limiter results ina load factor greater than or equal to 9.5g/-3.2g or ifsymmetric nong limiter maneuvering or asymmetric53

T.O. GR1F16CJ1
maneuvering exceeds a positive or negative g limit
specified in this section. Provide details of theoccurrence (maximum g indication, airspeed, alti
tude, description of maneuver, fuel weight anddistribution, etc.) and HUD videotape if it isavailable.
NOTEFSYM G limits apply to maneuvers

resulting from less than abrupt rollstick inputs and in which roll rate does
not exceed 20 degrees/second. ROLL G
limits apply to maneuvers resulting
from abrupt roll stick inputs or
maneuvers in which roll rate exceeds
20 degrees/second.

FA false maximum g indication may bedisplayed in the HUD due to INU
vibration while the aircraft is at
maximum g. G indications above 10(e.g., 0.2 for 10.2g) have been observed.

F DR Due to the location of the accelerometer, it should not be used todetermine maximum g force.

FFor evaluating a possible overg, determine the allowable CARRIAGE MAX
ACCEL G for the actual store configuration at the time of the occurrence. Use of
T.O. GR1F16CJ12, figure 59, NZW
curves (if applicable) is permitted.

F G's experienced during a wingtip vortex/

wake turbulence encounter should beconsidered as asymmetrical when determining if a g limit has been exceeded.

AOA AND ROLLING LIMITATIONS
Refer to figure 58. With heavy wing loadings, it may
be necessary to cancel the roll command up to 90
degrees early to avoid exceeding the maximum bankangle limit. Except for emergency conditions, do not
fly category III loadings with the STORES CONFIG
switch in CAT I.

FCategory III loadings are not protectedfrom AOA or rollinduced departureswith the STORES CONFIG switch in
CAT I.

54

FDamage to or failure of wing internal

structure can occur if rolling maneu
vers are performed with the STORESCONFIG switch in CAT I while
carrying a category III loading.

An asymmetric loading is any asymmetry that
requires roll and/or yaw trim. Refer to ASYMMET
RIC LOADINGS, Section VI.
Nose slice and yaw departure may occur during
maximum command rolls on the CAT I AOA limiterat high altitude when carrying a centerline tank.
Refer to YAW DEPARTURE, Section VI.

FIf the aircraft CG is near the aft limit,

departure may occur while performing
low airspeed, high AOA, maximum
command rolling maneuvers witheither of the following:

D Asymmetric category I missile loadings(station 2, 3, 7, or 8).

D Speedbrakes opened.

FThe indicated bank angle change limit

is particularly critical for category I
loadings with 370gallon fuel tanks
plus suspension equipment on stations
3 and 7. Care is required with theseloadings to check the roll so as not toexceed the indicated bank angle
change limit.

STORES LIMITATIONS
Refer to T.O. GR1F16CJ12, figure 510 for
authorized stores loading configurations and related
limitations.
ASYMMETRIC STORES LOADING
Refer to figure 317 for computation of asymmetricmoment. The maximum allowable asymmetric(rolling) moment for ground handling, taxi, takeoff,
inflight, and landing is 25,020 footpounds.
Takeoff is prohibited when the roll trim necessary tocompensate for an asymmetric loading exceeds the
maximum roll trim available. Refer to T.O.
GR1F16CJ11, PART 2.

T.O. GR1F16CJ1MISCELLANEOUS LIMITATIONSNET ARRESTMENT LIMITATIONS
Refer to figure 55 for cable/net arrestment limits.
61QSII/BAK15 multielement net barrier systems
have the following designations: 24A, 30F, 40A, HP,
and MEN. Any net barrier system with these
designations, e.g., 61QSII/HP30 or MK6/MEN, is
compatible with the aircraft.
The 61QSII/BAK15 arresting system consists of two
major components  an energy absorber and an
engagement device. The engagement device is the net
which is attached to the energy absorber. Theseabsorbers include heavy chains as used with theMA1A system, friction brake systems such as the
BAK9 and 12, and water twisters as used in BAK13
systems. Because of the variety of energy absorbers inuse, it is not possible to establish one maximum
engagement speed for the aircraft. In each case, themaximum engagement speed is determined by thecapabilities of the energy absorption system attachedto the net. For example, the HP30 is an all nylon,
three element net. It incorporates all the latest
technology and, coupled with a 1200foot runout BAK
12, should provide a 35,000pound F16 with190knot arrestment capability.
CROSSWIND LIMITS
Refer to T.O. GR1F16CJ11, Part 2 for crosswind
limits for takeoff and landing.

55

T.O. GR1F16CJ1

Instrument Markings
ENGINE F100-PW-229

800 C GROUND ENGINE START

97% MAXIMUM STEADY STATE

1070 C MAXIMUM STEADY STATE
FTIT

RPM

15 PSI MINIMUM

2850-3250 PSI NORMAL

95 PSI MAXIMUM

3250 PSI MAXIMUMOIL PRESSURE

HYDRAULIC PRESSURE

GR1F-16CJ-1-1129X37

Figure 51.(Sheet 1)56

T.O. GR1F16CJ1

Instrument Markings
ENGINE F110-GE-129

935 C ENGINE START

108% MAXIMUM

980 C MAXIMUM
FTIT

RPM

15 PSI MINIMUM

2850-3250 PSI NORMAL

65 PSI MAXIMUM

3250 PSI MAXIMUMOIL PRESSURE

HYDRAULIC PRESSURE

GR1F-16CJ-1-0129X37

Figure 51. (Sheet 2)

57

T.O. GR1F16CJ1Engine Limitations
ENGINE F100PW229
GROUND
FTIT
°C

RPM
%

OIL PSI

REMARKS

START

800

-

-

During cold start, oil pressure may be 100 psi for up to 1 minute

IDLE

625

-

15 (min)Maximum FTIT in SEC is 650°C

MIL/AB

1070

97

3095

At MIL and above, oil pressure must increase 15 psi minimum above
IDLE oil pressure. Use transient rpm limit for takeoff

TRANSIENT

1090

98

3095

Time above 1070° is limited to 10 seconds

FLUCTUA
TION

$10

$1

$5 IDLE

Must remain within steadystate limits. Inphase fluctuations of morethan one instrument or fluctuations accompanied by thrust surges indicate engine control problems.
problems Nozzle fluctuations limited to $2% at
and above MIL. Fluctuations not permitted below MIL

CONDITION

$10
above
IDLE

IN FLIGHT
FTIT
°C

RPM
%

OIL PSI

870

-

-

-

IDLE

-

-

15 (min)-

MIL/AB

1070

97

3095

Oil pressure must increase as rpm increases. Use transient rpm limit
with LG handle DN and for 3 minutes after LG handle is placed UP

TRANSIENT

1090

98

3095

Time above 1070° is limited to 10 seconds

FLUCTUA
TION

$10

$1

$5 IDLESame as ground operation. Zero oil pressure is allowable for periods up
to 1 minute during flight at less than +1g

CONDITION
AIRSTART

REMARKS

$10
above
IDLE

Figure 52.(Sheet 1)58

T.O. GR1F16CJ1Engine Limitations
ENGINE F110GE129
GROUND
FTIT
°C

RPM
%

OIL PSI

REMARKS

START

935

-

-

During cold start, oil pressure may be 100 psi for up to 2 minutes

IDLE

650

-

15 (min)-

MIL/AB

980

108

2565

At MIL and above, oil pressure must increase 10 PSI minimum above
IDLE oil pressure

TRANSIENT

980

109

2565

-

FLUCTUATION

$10

$1

$5

Must remain within steadystate limits. Nozzle fluctuations limited to$2

CONDITION

IN FLIGHT
FTIT
°C

RPM
%

OIL PSI

935

-

-

-

IDLE

-

-

15 (min)-

MIL/AB

980

108

2565

Oil pressure must increase as rpm increases

TRANSIENT

980

109

2565

-

FLUCTUATION

$10

$1

$5

Must remain within steadystate limits. Zero oil pressure is allowablefor periods up to 1 minute during flight at less than +1g

CONDITION
AIRSTART

REMARKS

Figure 52. (Sheet 2)

59

T.O. GR1F16CJ1

Operational Envelope

EngineENGINE F100-PW-229

NOTES:
Transfers from SEC to PRI should be performed below 40,000 feet MSL and subsonic with the throttle at
MIL or below.
Maximum operating airspeed is 800 knots from sea level to 30,000 feet MSL. Above 30,000 feet MSL, theaircraft is limited to 2.05 mach.

70

60

MINIMIZE THROTTLE
MOVEMENTS IN SEC

1000 FEET40

ALTITUDE

5030

DO NOT RETARD
THROTTLE BELOWMIL UNTIL SUBSONIC IN SEC

20UNRESTRICTED THROTTLE
MOVEMENT IN SEC

10

0
0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

2.0

2.2

MACH NUMBER

1F-16X-1-4018X

Figure 53.(Sheet 1)510

T.O. GR1F16CJ1EngineOperational Envelope
ENGINE F110-GE-129

70

60

ALTITUDE

1000 FEET

50

UNRESTRICTED THROTTLE
MOVEMENT IN SEC

40

DO NOT RETARD
THROTTLE BELOWMIL UNTIL SUBSONIC IN SEC

30

20

10

0
0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

2.0

2.2

MACH NUMBER

1F-16X-1-4019X

Figure 53. (Sheet 2)

511

T.O. GR1F16CJ1

Jet Fuel Starter Limits
CONDITION

Normal Ground Operation

LIMITS* Continuous motoring of the engine shall not exceed 4 minutes. After 4 minutes of con
tinuous motoring, a normal engine start may be initiated after 5 minutes of coolingA minimum wait of 1 minute is required after each JFS start attempt to allow fueldrainage from the JFS

Hot Start of Engine

Motor until FTIT is below 200°C

Airstart/In Flight

Below 20,000 feet MSL and 400 knots. 3minute maximum run time when the engine
is operating satisfactorily above 60 percent rpm; otherwise, unlimited

PW229
GE129
* Motoring is defined as JFS rotating the engine at
22 percent rpm,25
percent rpm
minimum with the throttle in OFF.

NOTES:D OAT between 20°F (-6°C) and 100°F (38°C). A minimum brake/JFS accumulator pressure of 3000 psi is required
for START 1 and 2800 psi is required for START 2.
D OAT between -25°F (-32°C) and 20°F (-6°C) or OAT above 100°F (38°C). START 2 and a minimum brake/JFS accu
mulator pressure of 2800 psi are required.
D OAT below -25°F (-32°C). START 2 and a minimum brake/JFS accumulator pressure of 3200 psi are required.
D If one brake JFS accumulator is depleted, verify a minimum pressure of 3000 psi in the remaining brake/JFS accu
mulator before attempting START 2.

Figure 54.

512

T.O. GR1F16CJ1

Cable/Net Arrestment LimitsCompatible Cable Systems With Established Aircraft Limits
BAK6
BAK9
BAK12 (Standard, Extended, and Dual)
BAK13 (Navy designation is E28)
BAK14
BAK15 (NI) (Net barrier with cable for hook)
MAAS*44B2L
ROUTINE ARRESTMENT
A planned event. Operational conditions are such that each landing requires arrestment. Such
operational conditions include operating from highways or from runways that are too short fornormal landings. The standard factor of safety is used to determine the maximum engagement
speed.

S PX II LIMITS - 150 knots for GW less than/equal to 42,300 pounds, 130 knots for GW greaterthan 42,300 pounds, (*156 knots).

S PX III LIMITS - 146 knots (*160 knots).
EMERGENCY ARRESTMENT
An unplanned event. A reduced factor of safety is accepted and the corresponding maximum en
gagement speed is slightly higher than for a routine arrestment. With a ground level ejection capa
bility, a reduced factor of safety for the arrestment is possible since the pilot can still eject if thearrestment fails.

S PX II LIMITS - 160 knots for GW less than/equal to 42,300 pounds, 140 knots for GW greaterthan 42,300 pounds, (*171 knots).

S PX III LIMITS - 160 knots (*174 knots).
Compatible Cable Systems Without Established Aircraft Limits

Use of arresting systems without established aircraft limits may resultin failure of the hook, hook backup structure, or nose landing gear.
BLISS 500S6
TAGS BLISS 500S6B Transportable
44B3H/SP/WRMAGS 44B3H/SP/WR Mobile
RHAG MK1
RHAG MK2
PUAG
PAAG
PIV/BAK12 Portable
Figure 55.(Sheet 1)

Change 1513

T.O. GR1F16CJ1

Cable/Net Arrestment LimitsCompatible Strut Engagement/Cable Systems Without Established Aircraft Limits

Use of arresting systems without established aircraft limits may resultin failure of the hook, hook backup structure, or nose landing gear.
MA1A Modified
MA1A/E5
MA1A/BAK9
MA1A/BAK12
NOTE
Attempting to engage an unmodified (nonhook capable) MA1A will most
likely be unsuccessful.

Compatible Net Barrier Systems
BAK15
SAFEBAR (Safeland Barrier)Stanchions
61QSII61QSIIMMK6
NOTE
The above stanchions can be rigged with the following nets: HP30, HP40,
MEN241, MEN301, 24, 30, or 40. In addition, they will be associated withthe following energy absorbers: Chain (E5), BAK9, BAK12, BAK13,
34B, 44B, 500S, BEFAB 6:3, BEFAB 21:2, and BEFAB 24:4.
Other Net Systems
ARZ30/40
BEFAB 21:2/MK VI1 stanchion
BEFAB 6:3/MK VI1 stanchion
MK12ARAF TYPE A
RAF TYPE B

Figure 55.(Sheet 2)

514

T.O. GR1F16CJ1

Airspeed Limitations(Systems)

Acceleration
Limitations
LOAD FACTOR (g)SYSTEM OR CONDITION

KIAS/MACHCONFIGURATION

70 (includes
ground windvelocity)

Canopy Open or in Transit

TAKEOFF

LG E
Extendedt d d or iin T
Transit
it

300/0.65, which
ever is less

AR Door Opening/Closing

400/0.85, which
ever is less

AR Door Open

400/0.95, which
ever is less

Flight in Severe Turbulence
(+3g)

500

Drag Chute Deployment

170

SYMMETRIC ASYMMETRIC+4 0 00+4.0,
0.0+2 0 00+2.0,
0.0+2 0 00+2.0,
0.0+2 0 00+2.0,
0.0

LANDINGLG RETRACTION*

LG EXTENSION* If the LG handle is raised near 2 g's approaching 300 knots, actua
tor power may be insufficient to completely retract the LG until g isreduced.

Figure 56.

Figure 57.

AOA and Rolling Limitations
LOADING
CATEGORYSTORESCONFIGSWITCHMAX AOAI

I

LIMITER

360°

III

III

LIMITER360°

MAX BANK ANGLE CHANGE FOR MAX ROLL MANEUVERNOTES:
1.

Determine loading category from the appropriate line in T.O. GR1F16CJ12, figure 510, Stores Limita
tions.

2.

The roll command should be released in sufficient time to avoid overshooting the indicated bank angle change
limits.
Figure 58.

Change 1515/(516 blank)

T.O. GR1F16CJ1

SECTION VI
FLIGHT CHARACTERISTICS
TABLE OF CONTENTS
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Flight Control System . . . . . . . . . . . . . . . . . . . . .
FLCS Limiters . . . . . . . . . . . . . . . . . . . . . . . . . .
Leading Edge Flaps . . . . . . . . . . . . . . . . . . . . .
Speedbrakes . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Autopilot . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Trim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Normal Flight Characteristics . . . . . . . . . . . . . .
Category I Loadings . . . . . . . . . . . . . . . . . . . . .
Category III Loadings . . . . . . . . . . . . . . . . . . .
Conformal Fuel Tanks PX III . . . . . . . . . . . . .
Flight With LG Down . . . . . . . . . . . . . . . . . . . .
Landing Configuration . . . . . . . . . . . . . . . . . .
Factors Affecting Flying Characteristics . . . . .
CenterofGravity Considerations . . . . . . . . .
Effect of Thrust . . . . . . . . . . . . . . . . . . . . . . . . .
Effect of Low Airspeed Maneuvering . . . . . .
High Pitch, Low Airspeed . . . . . . . . . . . . . . . .
Flight With Stores . . . . . . . . . . . . . . . . . . . . . .
Limit Cycle Oscillation andAeroservoelastic Oscillation . . . . . . . . . . . .
Asymmetric Loadings . . . . . . . . . . . . . . . . . . .
Store Separation . . . . . . . . . . . . . . . . . . . . . . . .
OutofControl Characteristics . . . . . . . . . . . . . .
Yaw Departure . . . . . . . . . . . . . . . . . . . . . . . . . .
Pitch Departure . . . . . . . . . . . . . . . . . . . . . . . . .
Deep Stall . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Spin . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Recoveries . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
SelfRecovery . . . . . . . . . . . . . . . . . . . . . . . . .
Deep Stall Recovery . . . . . . . . . . . . . . . . . . .
Spin Recovery . . . . . . . . . . . . . . . . . . . . . . . .
Engine Operation During Departures/OutofControl . . . . . . . . . . . . . . . . . . . . . . . .
Degraded Flight Controls . . . . . . . . . . . . . . . . . .
FLCS DBU . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Leading Edge Flaps Locked(Symmetric) . . . . . . . . . . . . . . . . . . . . . . . . . .
Standby Gains . . . . . . . . . . . . . . . . . . . . . . . . . .
One Hydraulic System . . . . . . . . . . . . . . . . . . .
Speedbrakes . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Aircraft Damage . . . . . . . . . . . . . . . . . . . . . . . . . . .
Horizontal Tail . . . . . . . . . . . . . . . . . . . . . . . . . .
Flaperon . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Rudder and Ventral Fins . . . . . . . . . . . . . . . . .

61
61
61
61
62
62
62
63
63
63
63
63
64
64
65
65
65
66
66
67
67
68
68
68
69
610
610
611
611
611
612
613
613
613
613
613
613
613
614
614
614
614

Leading Edge Flaps . . . . . . . . . . . . . . . . . . . . .
Wing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Radome . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Dive Recovery . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

614
614
614
615

INTRODUCTION
Information presented in this section reflects theflight characteristics with category I and III
loadings. Refer to AOA AND ROLLING LIMITATIONS, Section V, for information regarding specificcategories.

FLIGHT CONTROL SYSTEMThe FLCS is a fourchannel flybywire system. The
FLCC combines pilot commands along with aircraft
motion and flight conditions to command position of
the flight control surfaces. Artificial stability
provided by the FLCS allows for relaxed static
stability which increases performance and maneu
verability by reducing trim drag and increasing
maximum lift. Refer to FLIGHT CONTROLSYSTEM, Section I.
FLCS LIMITERS
FLCS limiters may be defeated if maneuvering
limits are not strictly observed. Departure may
result from maximum maneuvering combined withmaximum permissible aft CG. The most criticalmaneuvers are maximum command rolls coupledwith either maximum aft stick or exceeding the
maximum bank angle change limits.
The AOA/g limiter depends on the horizontal tails
to control g and AOA. If the airspeed decreases until
there is not enough airflow over the tails to provide
this control, the limiter is defeated and a departureor deep stall may result. This condition may occur
in a nosehigh, decreasing speed maneuver. Refer to
LOW AIRSPEED OPERATING LIMITATIONS,
Section V.
LEADING EDGE FLAPSThe LEF system is designed to optimize wing airflow.
It also provides special functions in the takeoff andlanding configurations.
61

T.O. GR1F16CJ1
At subsonic speeds, the LEF's move from 2 degrees
up to 25 degrees down as a function of mach number,
AOA, and altitude. This automatic operation
significantly reduces buffet and drag and improves
high AOA directional stability. If the LEF's fail toschedule properly during maneuvering flight,
higher than normal buffet levels occur and, in thehigh AOA region, reduced directional and longitu
dinal stability may also be noted. At supersonic
speeds, the LEF's are scheduled to minimize drag.
SPEEDBRAKESThe speedbrakes provide deceleration over theentire flight envelope. There are no trim changes
associated with speedbrake operation and induced
buffet is negligible. A yaw oscillation may occur at
approximately 1.4 mach with speedbrakes opened.
The oscillation is neutrally damped and no action isrequired. The oscillation may be eliminated byeither closing the speedbrakes, reducing mach, orincreasing the g level.
AUTOPILOT
With the HDG SEL and ALT HOLD modes engaged,
the aircraft turns, climbs, or dives within the limitsof the autopilot to capture the heading reference
and the altitude reference regardless of aircraft
attitude. This autopilotcommanded flight may
eventually return the aircraft to a preselectedheading and altitude if airspeed and altitude
permit.
If the ALT HOLD remains engaged as airspeedtransits 1.0 mach, a mild pitch transient may occur
and can be eliminated by depressing the paddle
switch until the altimeter has stabilized.
Use of pitch altitude or attitude hold during
decelerating flight can produce either autopilot
disengagement and the FLCS A/P FAIL PFL or descentfrom the referenced altitude if AOA increases abovecertain values. In CAT I, the autopilot disengages andthe FLCS A/P FAIL PFL occurs when 15 degrees AOA
is exceeded. In CAT III, the aircraft starts descending
from the referenced altitude at approximately 810degrees AOA. If the descent is not corrected within 5
seconds, the FLCS A/P DEGR PFL occurs.

Since the autopilot command is additive
to stick commands, use of ALT HOLD in
conjunction with high g maneuvering
may result in aircraft overg.
62TRIM
The aircraft can be trimmed about all three axes.
With pitch trim centered in cruise gains and no
input to the stick, the aircraft attempts to maintain1g flight regardless of flight condition unless AOA
exceeds 15 degrees. Full noseup/full nosedown trim
corresponds to +3.4g or -1.4g in cruise gains.
NOTE
Airspeed must be closely monitored
because there is little aerodynamic
indication of large changes in airspeed.
Cues which normally indicate airspeed
changes, such as stick movement or
trim changes, are absent.
Above 15 degrees, the FLCS commands anincreasing nosedown pitch attitude as a warning ofdecreasing airspeed. A specific force applied to thestick commands a specific g increment from the trim
condition. Moving the PITCH TRIM wheel changes
the handsoff trim condition.
In takeoff and landing gains, zero pitch trim
commands zero pitch rate until 10 degrees AOA. A
slight amount of noseup trim is required to zerostick forces during an 1113 degree AOA approach.
When properly trimmed and no input is applied tothe stick, the aircraft attempts to maintain zero rollrate. Moving the ROLL TRIM wheel changes thehandsoff trim condition. Maximum roll trim
authority is approximately onefifth of maximum
stick command of cruise gains. However, precise
trimming is difficult using the stick TRIM button.
Roll trim requirements may change with stores,
particularly at supersonic speeds. For asymmetricconfigurations (asymmetrical stores or rudder
mistrim), roll retrimming may be required asflight conditions change. Roll trim inputs alsocommand rudder deflection through the ARI. The
ARI switches out with wheel spinup upon landing.
Likewise, the ARI switches in following takeoff as
the wheels spin down. This switching may cause
abrupt rudder inputs to occur if roll (due toasymmetries or crosswind) is being input via the
stick or trim.
Rudder trim inputs command rudder deflection.
Rudder trim is required with asymmetricalconfigurations and frequently during supersonicflight, especially with stores. Maximum trim
authority is 12 degrees.

T.O. GR1F16CJ1NORMAL FLIGHT CHARACTERISTICS

The capability of the aircraft to rapidly
attain and sustain high g levels, which
may cause ginduced loss of conscious
ness, should be considered during
heavy maneuvering.
The FLCS provides constant response for specificinputs regardless of flight conditions. Commanded
pitch responses are in g increments per stick force
for AOA below 15 degrees. Above 15 degrees AOA,
stick force increases as a cue of increasing AOA.
Conventional cues such as aircraft buffeting forces
are not always present as AOA and g limits areapproached. The commanded lateral response isroll rate per stick force. Rudder position iscommanded by rudder pedal force.
The ARI provides coordinated rudder commands
and reduces sideslip during rolling maneuvers.
Additional pilot rudder commands do not improveroll performance but do increase departure
susceptibility. When ARI is not available during
takeoff and landing (MLG wheel speed above 60
knots), pilot rudder commands may be required toprovide coordinated flight and to control yaw.

Rolling g limits are not protected by
the FLCS and must be observed.
CATEGORY I LOADINGSCATEGORY III LOADINGS
Aircraft response with most category III loadings
remains similar to that of the clean aircraft;however, large stores significantly increase total
aircraft drag and reduce performance. Lightbuffeting may occur during level flight at approxi
mately 0.92 mach. In addition, surging may occur
near the store limit airspeed, especially at low
altitude. Neither condition requires specific action.
With STORES CONFIG switch in CAT III, theAOA/g limiter provides departure resistance for all
category loadings. Except for the requirement toavoid structural overstress, pilot workload is
reduced to a level comparable to that with category
I loadings.
CONFORMAL FUEL TANKS PX III
In general, the presence of CFT's only minimallyaffects flight characteristics. One item of signifi
cance is a reduction in directional stability which
manifests itself as higher angles of sideslip duringlateraldirectional maneuvering. Directional looseness" may be evident during tasks such as tracking
and aerial refueling.
LESS b6n Significantly higher angles of sideslip can begenerated with pilot rudder input. Rudder should be
used only as required. Avoid abrupt and/or large
inputs.

LESS b6n Maximum command and/or
abrupt pilot rudder inputs can result in
departure or structural overload when
CFT's are installed.

The FLCS minimizes the possibility of departures or
spins. Roll rate inputs command flaperons andhorizontal tails for roll power to provide a relatively
constant roll response.

FLIGHT WITH LG DOWN

Maximum command 360degree rolls at subsonic
speeds may cause a slight g reduction on
termination. At supersonic speeds, maximum rollrates may cause a slight increase in g. At high AOA
and low airspeed conditions, roll performance isreduced by the FLCS to minimize pitch/roll coupling.
Aft CG's, open speedbrakes, asymmetric missiles, orcenterline stores decrease departure resistance.

With the LG handle down, LG and TEF's areextended and the FLCS operates in takeoff andlanding gains. Normally, this mode of flight islimited to takeoff, approach, and landing; however,circumstances can arise which require flight for an
extended distance with the LG down. If so, the LGshould be left pinned but the streamers should be
removed to prevent damage.

63

T.O. GR1F16CJ1
With the LG pins installed, it is preferable to raise
the LG handle once airborne. This action retracts the
TEF's and significantly reduces drag and the FLCS
switches to cruise gains. For cruise with only the LG
down, the best airspeed is 230250 knots. A clean
aircraft can be flown at 25,00030,000 feet with the
LG down and TEF's up and fuel flow is 30003400
pph. If the LG handle is left down, the TEF's remain
down and the best cruise altitude is less than 20,000
feet with significantly higher fuel flows.

All normal landings should be made with speedbrakes opened to the 43degree position to avoid afloating tendency when entering ground effect. A
touchdown at the desired point at 13 degrees AOA canbe achieved when flying final at either 11 or 13
degrees AOA by adjusting the initial aimpoint.

LANDING CONFIGURATION

Due to the aircraft light wing loading and thefloating tendency associated with ground effect,
wake turbulence on final approach and during
touchdown presents a significant hazard.
Increased spacing between landing aircraft should
be used when there is little or no effective
crosswind. Exercise caution and be ready to initiate
a goaround when wake turbulence is encountered.
An early goaround decision may help avoid theneed for a large roll control input. Such an input
retracts a flaperon, causing decreased lift andpossibly a sink rate as well as a roll. A large roll
input at slow airspeed also causes a large horizontal
tail split. A horizontal tail surface could contact the
runway while trying to counter wake turbulence
effects during touchdown.

Two distinct techniques may be used when landing.
One technique is to trim for approximately 11
degrees AOA and to fly that airspeed throughout
the final approach. Attitude/glidepath is controlled
by the stick, and airspeed/AOA is controlled by thethrottle. This technique allows better pitch control,
better overthenose visibility, and a more stable
HUD presentation. In gusty wind conditions, theaircraft wallows less, and during the flare, the sinkrate is easier to control. The aircraft floats
approximately 8001200 feet from flare initiation totouchdown. Another technique is to trim for 13
degrees AOA and to fly that airspeed throughout
the final approach. The throttle is used primarily tocontrol glidepath, and the stick controls airspeed
through control of AOA and direction through bankangle. This type of approach primarily allows better
control of touchdown point and more efficientenergy dissipation; however, since the aircraft isalready at 13 degrees AOA, the flare is moredifficult, and care must be exercised to avoid
scraping the speedbrakes or landing firm. The
aircraft floats approximately 500700 feet fromflare initiation to touchdown.
Regardless of the technique used, establish
computed final approach airspeed for the desired
AOA early on final and trim the aircraft. Airspeed
changes result in pitch changes, which may require
retrimming and make glidepath control moredifficult. PW 229 Small throttle adjustments may berequired as the DEEC retrims the engine.
On short final, avoid premature or large thrust
reductions which may cause increased sink rates anda hard landing. Use thrust rather than back stick tocontrol undesirable sink rates. Increased back stick
may result in a tail strike in this situation. AOA
decreases slightly as the aircraft enters ground effect.

64

Increased control inputs to achieve normal aircraft
response as airspeed decreases are unnecessary.
Control inputs should be kept small to avoid
overcontrol.

If pitch trim is used during the turn to final, forward
stick/trim will be required upon rollout on final
approach to counter noseup motion. Floating
tendencies following a high flare or aircraft bounce
may be increased. Slight forward stick force may berequired to prevent a long or slow landing. Stickforce per degree AOA change is reduced and should
not be relied upon as a slow speed cue.

FACTORS AFFECTING FLYING CHARACTERISTICS
NOTE
Pitch sensitivity and pilot induced
oscillations (a maximum of " 0.5g)
may occur above 0.80 mach when
flying with 600gallon fuel tanks. Thisbehavior can be minimized by avoiding
large pitchstick inputs and rapid
pitchstick reversal. If this behavior
becomes objectionable, reduce airspeed and pitchstick inputs.

T.O. GR1F16CJ1
NOTE
FMomentary

uncommandedpitchchanges (a maximum increase of 1g)
and/or bank angle changes (a maxi
mum increase of 15 degrees) may occur
above 0.85 mach when flying with
600gallon fuel tanks. This behavior is
known to occur with loaded TER racks
but may occur with other store
loadings. Avoid overresponding to thechanges and use smooth stick input tominimize pilot induced oscillations.
Airspeed should be reduced if thisbehavior becomes objectionable. If
flying within normal load factor
carriage (MAX ACCEL G) limits
defined in T.O. GR1F16CJ12,
STORES LIMITATIONS, an incre
mental 1g uncommanded pitch change
will not exceed structural limits.

FA mild pitch oscillation (a maximum of"0.15g at 3 cycles per second) may
occur at 0.750.90 mach while in cruise
gains or at 330400 knots while intakeoff and landing gains. The oscillation is caused by the normal responseof the aircraft and FLCS and does notcause a significant tracking problem.

FMomentary roll hesitations may occur

during elevatedg maneuvering on ornear the AOA/g limiter with certain
loadings. This behavior is most notice
able between 250 and 350 knots when
flying above 25,000 feet MSL.

CENTEROFGRAVITY CONSIDERATIONSMonitoring the forward and aft fuel distribution
provides an indication of the aircraft CG.
As CG moves aft, higher pitch rates are obtainable
and susceptibility to departure and deep stall
increases.
NOTE
F C

The most aft CG occurs withapproximately 2000 pounds of internalfuel remaining.

F D With external fuel tanks, the mostaft CG occurs when the external fueltanks have just emptied.

EFFECT OF THRUST
Thrust changes result in little or no change inaircraft trim or stability at all operational loadfactors and for all store loadings.
EFFECT OF LOW AIRSPEED MANEUVERING

FMomentaryroll hesitations may
occur when commanding low tomoderate roll rates (generally lessthan 100 degrees per second) whenairspeed is above 350 knots, andaltitude is 20,000 feet or less. Thisbehavior is most noticeable when
flying without stores.

FA short duration (less than 3 seconds)

series of rapid wing rocks (less than
10 degrees of bank angle change) may
occur when terminating a high rollrate maneuver at airspeeds above 400
knots. This behavior is most notice
able when flying without stores.

FMinor AOA oscillations (less than "2

degrees) may be noticed during elevatedg maneuvering on or near the
AOA/g limiter with certain loadings.
This behavior is most noticeable
between 250 and 350 knots when
flying above 25,000 feet MSL.

Departures are possible at low airspeeds and low
pitch angles if large, simultaneous pitch and roll
inputs are made.
The FLCS requires adequate airflow over thecontrol surfaces to be effective, which means thatairspeed is a critical factor in departure susceptibil
ity during maneuvering. Low airspeeds should,
therefore, be avoided during maximum perfor
mance maneuvering.

FFLCS limiters can be defeated at low

airspeeds (below 200 knots in a CAT I
configuration) during maximum pitch
and roll commands initiated frombelow limiter AOA's.

F The aircraft can be departed (fromparameters outside the tone on area of
figure 142) with no low airspeed warning
tone present, if abrupt or uncoordinated
FLCS commands are made.

65

T.O. GR1F16CJ1
HIGH PITCH, LOW AIRSPEED
The low airspeed warning tone sounds to aid inrecognizing that critical high pitch, low airspeed
flight conditions are reached.

Proper assessment of flight path angle(not pitch angle) is key to determining
the nearest horizon and performing aproper recovery. Differences between
flight path and pitch angle of up to 25
degrees, combined with the visual
illusion caused by a reclined seat canlead to an incorrect decision tocontinue the maneuver through thevertical. The risk of a departure/deepstall in this instance is very high.
Avoiding a departure under these conditions requires
specific control techniques. To recover, first release
aft stick pressure. This action unloads the aircraft
and reduces AOA so that the flightpath more closely
coincides with the longitudinal axis of the aircraft.
Smoothly roll inverted to the nearest horizon. Afterthe roll, smoothly apply the aft stick pressure
required to keep the nose moving toward the horizon.
As airspeed continues to decrease during therecovery, more aft stick pressure may be required tokeep the nose moving. Continue to smoothly increase
aft stick pressure up to the AOA/g limiter. If full aftstick is inadvertently released, do not reapply it
unless required to keep the nose moving.

Avoid large, simultaneous pitch androll commands to preclude a rollcoupled departure. Small lateral commands can be made as required tomaintain wings level, inverted flight.
Do not abruptly apply aft stick pressure
at anytime during the recovery. Rapid
aft stick pressure will generate exces
sive AOA, overshooting the AOA
limiter and causing departure.
During a recovery where full aft stick is required,
nose movement toward the horizon may slow down
markedly as the AOA/g limiter tries to limit AOA.
As long as the nose continues to move, no further
action is required. If the nose of the aircraft does notcontinue to move toward the horizon, the aircraft
has departed, and outofcontrol recovery procedures should be initiated.
66

After attaining a nosedown attitude with airspeed
increasing, continue to avoid abrupt commands.
The aircraft may either be unloaded and rolled
upright or a splits recovery can be made at airspeed
above 200 knots, altitude permitting, before
continuing to maneuver. The splits recovery is thesimplest way to recover the aircraft. However, if
altitude is a factor, allow airspeed to increase to aminimum of 150 knots, unload the aircraft to lessthan 1g, smoothly roll upright, and recover to level
flight.

FLIGHT WITH STORES
The major effects of stores are increased weight andinertia. A reduction in aircraft response anddamping should be expected as GW increases,
particularly when stores are carried. Stores
generally reduce longitudinal and directional
stability and increase inertial effects so that thepilot must anticipate initiation and termination of
maneuvers based on the loadings. High roll andpitch rates are attainable with full force applicationof the stick. Avoid abrupt control commands which
may cause AOA overshoots in excess of thelimitations specified in Section V and T.O.
GR1F16CJ12.
Bank angle change limits must not be exceeded.
During rolling maneuvers with category III
loadings, the roll rate must be stopped prior to360degree bank angle change. Removing the roll
input is not always sufficient (opposite stick may berequired). Refer to STORES LIMITATIONS, T.O.
GR1F16CJ12, for carriage limits.
Certain store loadings may exhibit decreased
yaw/roll damping in supersonic flight and result in
mild yawing oscillations. Neutral and divergent
yaw and roll oscillations may occur during sideslip
maneuvers at supersonic airspeed. These oscilla
tions are aggravated when large stores are carried.
Excessive vertical tail loads may be generated if
oscillations become sufficiently large. If oscillations
are encountered during rudder commands, release
the rudder input. Additionally, buffeting may occur
in transonic flight with certain store loadings.
NOTE
A mild airframe vibration may be
experienced while supersonic whencarrying a centerline store.

T.O. GR1F16CJ1
LIMIT CYCLE OSCILLATION AND AEROSERVO
ELASTIC OSCILLATION
A limited amplitude constant frequency oscillation(commonly referred to as limit cycle oscillation or
LCO) may occur with certain stores loadings. The
LCO (typically 510 cycles per second) may occur inlevel flight or during elevated g maneuvers. The
LCO may appear as buffeting or turbulence similar
to that experienced during normal transonic buffet,
but the buffeting is a constant frequency, lateral
acceleration from sidetoside or, in some cases,
vertical accelerations up and down. The magnitude
generally increases with increasing airspeed and/or
load factor. Other cues of LCO include significant
vertical movement of the forward area of wing
stores, especially wingtip launchers and missiles;
this motion is typically up and down, but may also
follow a circular pattern. In addition, cockpit
instruments may become difficult to read as the
LCO amplitude increases from moderate to severe.
Within published carriage limits, LCO is not
detrimental to the aircraft. LCO susceptible
loadings include airtosurface and airtoair
loadings and associated downloadings. If LCO is
encountered and is uncomfortable or distracting,
reduce airspeed and/or load factor. Refer to
STORES LIMITATIONS, T.O. GR1F16CJ12, forcarriage limits.
An aeroservoelastic (ASE) oscillation is similar to
LCO. Wing and store oscillation and cockpit
vibration may be indistinguishable from thosecaused by LCO. However, ASE oscillation is driven
by the FLCS, resulting in key differences. ASE
oscillation (typically at 45 cycles per second) ismost likely to occur within a narrow range between
0.9 and 0.95 mach. The magnitude is strongly
dependent on mach, but not strongly dependent onload factor, and increases in severity as altitudedecreases. ASE oscillation will probably occur whencarrying wingtip AIM120 missiles. The presence ofstores at stations 3 and/or 7 may dampen theoscillation. Within published carriage limits, ASE
oscillation is not detrimental to the aircraft. If ASE
oscillation is encountered and is uncomfortable or
distracting, change airspeed by at least 0.05 mach.
NOTE
LCO and ASE oscillation may be
indistinguishable to the pilot. Either
may produce severe oscillation at themost critical flight condition. While
not detrimental to the aircraft withinpublished carriage limits, the motions
may be extremely uncomfortable or

impact mission accomplishment. The
most effective way to reduce LCO or
ASE is to reduce airspeed.
ASYMMETRIC LOADINGS
If roll trim is used to hold up a heavy wing, the ARI
adds rudder in the direction of the roll trim, causinga yaw away from the heavy wing. If roll trim is used
for takeoff, yaw occurs when the wheel speed drops
below 60 knots groundspeed after takeoff, activat
ing the ARI. This yaw is easily controllable byrudder commands. Yaw and roll trim requirementschange for different flight conditions.
Asymmetric loads increase departure and spin
susceptibility. Roll commands/trim away from theheavy wing is required to maintain the desired roll
attitude. Increasing g requires additional rollcommands/trim. Therefore, aft stick commands
result in increased roll requirements which, inturn, produce yaw away from the heavy wing due to
ARI action.

FWith certain asymmetric category III

loadings (2000 pounds or greater on
station 3 or 7 with stores on stations 4,6, and/or 5), rapid or abrupt aft stickcommands may result in sudden noseslicing departures.

FDeparture with an asymmetric category III loading may result in a fast,
flat (possibly nonrecoverable) spin.
NOTE
Leftwing heavy asymmetries are
more susceptible to departure.

During TF, commanded flyups with asymmetricloads result in a slower roll to wings level away fromthe heavy wing. Stick inputs to assist the roll to wings
level may be required as described in T.O.
GR1F16CJ3411
At high airspeeds, asymmetric loads exhibit some
unusual flight characteristics. Frequent trim
reversals may occur during supersonic acceleration. At airspeeds greater than 700 knots, yaw
oscillation may occur with significant lateral
accelerations.
Over 750 knots, a high frequency directional
shaking may occur with loadings such as the ECM
pod.
67

T.O. GR1F16CJ1STORE SEPARATIONYAW DEPARTURE

Symmetrical store releases and wingtip AIM9
missile launches can be accomplished with nounusual aircraft responses. Separation of the
300gallon fuel tank produces negative g on theaircraft. The magnitude of this response depends on
the amount of fuel remaining in the tank and themach number at release. Separating a full centerline
fuel tank at supersonic speeds produces the worst
response (up to -1.5g).

A yaw departure occurs when sideslip increases
beyond the normal controllable range (i.e., beyond
about 15 degrees). The primary indication of mostyaw departures is an abrupt nose slice. The aircraft
then fails to respond properly to pilot commands
and exhibits uncommanded motions. AOA is in thenormal range (-5 to +25 degrees) during the initial
phase of the nose slice. Immediately following a yaw
departure, a pitch departure usually occurs
resulting in AOA indication of -5 or +32.

Separation of 370gallon fuel tanks produces aminimal aircraft response. Separation of a single
370gallon fuel tank will initially produce aircraft
positive g response and roll away from the separated
tank (up to +1g and 15 degrees of bank).

OUTOFCONTROL CHARACTERISTICS
A departure is a loss of aircraft control that ischaracterized primarily by uncommanded aircraft
motions or failure of the aircraft to respond to control
commands. In a pitch departure, the AOA increases
beyond the normal controllable range. In a yaw
departure, the sideslip angle increases beyond thenormal controllable range first, although a pitch
departure may immediately follow. The automaticfeatures of the FLCS normally prevent departures.
However, departures may occur when Section V
limits are exceeded or in certain circumstances when
the FLCS provides only marginal protection.

It is possible for the sideslip to briefly exceed thenormal controllable range without the aircraft
experiencing uncommanded motions. In this situation, the pilot's only indication of a departure may be
noticeable sideforces. These brief departures typical
ly selfrecover within 5 seconds.
A yaw departure may occur while maneuvering on ornear the CAT I AOA limiter in the 0.80 to 0.95 mach
range, especially at high altitude (above 25,000 feet).
Maneuvering at high altitude is more critical than
low altitude because mach effects reduce directional
stability. These yaw departures usually result frommaximum command left rolls; but they may also
occur during symmetric maneuvering.

D 300gallon fuel tank.

The possibility of a yaw departure is increased
whenever the aircraft is configured with stores orsuspension equipment, especially a centerline store.
In general, the possibility of departure increases as
the number, weight, and size of such equipment orstores increases. Susceptibility to a yaw departure
increases significantly with lateral asymmetry, withleftwingheavy loadings being more likely to departthan rightwingheavy loadings. In addition, aheavier GW aircraft is generally more likely toexperience yaw departures. CAT I loadings mostsusceptible to yaw departures have one or more of thefollowing characteristics:

D 370gallon fuel tanks.

D Centerline store.

D Inlet mounted pod(s).

D Inlet mounted pod(s).

D Combination of a centerline store plus stores (orD Lateral asymmetry greater than 300 pounds at

With these loadings, maneuvering at high altitude(above 25,000 feet) increases the probability of a yaw
departure. Even moderate control commands may
cause aircraft with some CAT I loadings to departabove 25,000 feet.

Yaw departures can be minimized by avoiding abrupt
maneuvers in the 0.80 to 0.95 mach range with acenterline store, especially above 25,000 feet. Either
unload the aircraft prior to making roll commands orcommand only minimum required roll rate when
operating near the CAT I AOA limiter. Withcenterline store loadings having lateral asymmetries
greater than 300 pounds and inlet mounted pod(s),avoid abrupt aft stick commands above 25,000 feet.

With a lateral asymmetry in excess of 300 pounds(including wing tip missile and internal/externalfuel), abrupt maneuvering on or near the CAT I AOA
limiter can result in a departure if the aircraft isconfigured with any of the following:suspension equipment) at stations 3 and/or 7.

Abrupt maneuvering on or near the CAT I AOA
limiter at slow airspeeds (less than 200 knots) may
result in a departure.
68

stations 1, 2, or 3.

T.O. GR1F16CJ1

The probability of a yaw departuresignificantly increases above 25,000
feet for CAT I loadings having acenterline store, a lateral asymmetry
greater than 300 pounds, and inlet
mounted pod(s). With these loadings,
moderate, fullaftstick inputs at
35,000 feet and 300 knots have caused
yaw departures and spins.
A yaw departure may also occur with large
airtosurface lateral asymmetries with the STORESCONFIG switch in the CAT III position. These
departures can be avoided if abrupt control
commands are not used with lateral asymmetries inexcess of 1500 pounds at station 3 (or equivalent).
A yaw departure results in one of the following:
D A selfrecovery. The selfrecovery may occur quickly(within approximately 5 seconds). If a pitch
departure follows the yaw departure, the selfrecov
ery may require 1020 seconds. Random andpossibly abrupt pitch, roll, and yaw rates may occur.

D A deep stall.
D An upright spin.

PITCH DEPARTURE
A pitch departure occurs when the AOA exceeds the
AOA/g limiter. A pitch departure is classified either
as upright if the AOA is positive or as inverted if the
AOA is negative. Although the AOA indicator
displays a maximum of 32 degrees and a minimum of-5 degrees, the actual AOA during a departure will
exceed these values. In highly oscillatory departures,
the AOA indicator may momentarily indicate an AOA
below 32 degrees. Airspeed indications are erroneousand generally oscillate between the minimum value
and approximately 150 knots.
An upright pitch departure occurs when the AOA
exceeds the positive AOA/g limiter. Above 25 degrees
AOA, both horizontal tails are commanded to fulltrailing edge down by the pitch axis of the FLCS to try
to reduce the AOA. If AOA exceeds 35 degrees, theyaw rate limiter provides antispin commands to therudder, flaperons, and horizontal tails. Pilot roll andrudder commands are inhibited and pitch stickcommands are ineffective without use of the MPO
switch.
An inverted pitch departure occurs at negative AOAs
when the AOA significantly exceeds the negative glimiter. During the departure, the pitch axis of the

FLCS commands the horizontal tails to full trailingedge up to try to return AOA to the normal range.
Pitch stick commands are ineffective without use of
the MPO switch.
If AOA is below -5 degrees and airspeed is less than
170 knots, the yaw rate limiter provides antispincommands to the rudder. During an inverted
departure, roll and rudder commands should be
avoided. Pilot roll and rudder commands are
inhibited when the MPO switch is engaged.
An upright or inverted pitch departure can occur
when the aircraft is flown to airspeeds below thatindicated by the low speed warning tone. Inverted
pitch departures usually result from inverted flightat high pitch attitudes and low airspeeds, such as
those often encountered by going over the top at too
slow an airspeed. An upright or inverted pitch
departure may also occur at any pitch attitude if
abrupt stick commands are made at airspeeds below
200 knots. Simultaneous abrupt roll and aft stickcommands are especially likely to cause an upright
pitch departure. However, the aircraft can be safely
flown to the AOA/g and roll limiters below 200 knots
with smooth commands.
An upright pitch departure may also result from ayaw departure. In addition, recovery from aninverted pitch departure may cause the aircraft topendulum into an upright departure and vice versa.
The likelihood of a pitch departure increases if theaircraft is configured with stores (especially 370gal
lon fuel tanks), if the speedbrakes are opened, or if the
CG is near the aft limit.
Pitch departure characteristics are strongly in
fluenced by the airspeed and aircraft rates present at
departure. A low airspeed departure (below 200
knots) may have relatively benign uncommanded
pitch, roll, and yaw motions. Higher airspeed
departures are usually very dramatic with large
uncommanded pitch, roll, and yaw motions which
may persist for 10 seconds or more.
It is possible for the AOA to briefly exceed the AOA/glimiter
withouttheaircraftexperiencing
uncommanded motions. In this situation, the pilot's
only indication of a departure may be a failure of theaircraft to respond to control commands. These brief
departures typically selfrecover within 5 seconds.
A pitch departure results in one of the following:
D A selfrecovery which occurs within 520 seconds.

Random and possibly abrupt pitch, roll, and yaw
rates may occur.

D A deep stall.
D A spin.

69

T.O. GR1F16CJ1DEEP STALL
If the aircraft does not selfrecover following adeparture, a deep stall may have developed. A deep
stall is an outofcontrol flight condition in which theaircraft stabilizes at an AOA of approximately 60
degrees (upright) or -60 degrees (inverted) with low
yaw rates. The FLCS attempts to return AOA to thenormal range by commanding full horizontal taildeflection. However, the full horizontal tail deflection
is insufficient to return AOA to the controllable
range. The aircraft has entered a deep stall if the AOA
remains outside the controllable range. In a deep stall
the AOA indicator will be pegged at 32 or -5 degrees.
Recovery to controlled flight requires that the pilot
pitch rock the aircraft with the MPO switch in OVRD.
The MPO switch allows the pilot to override the FLCS
and to manually control the horizontal tails.
Airspeed indications are erroneous in a deep stall andfluctuate between 0150 knots. Altimeter indications
should be considered reliable; however, aircraft
oscillations may cause momentary stabilized or evenslightly increased altitude indications. Sink rate in a
deep stall is usually between 10,000 and 15,000 feetper minute. The normal load factor is approximately
1g or -1g for upright and inverted deep stalls,
respectively.
Upright deep stalls may be very stable with little orno pitch, roll, and yaw motions or may be very
oscillatory with large pitch, roll, and yaw motions.
Generally, a clean configuration results in a deep stallwith a near wingslevel pitching motion. If stores arebeing carried, especially a 300gallon fuel tank or
370gallon fuel tanks, the deep stall may be very
oscillatory, masking the pitch motions.
In an upright deep stall, the nose of the aircraft
usually oscillates "15 degrees about a slightly
noselow pitch attitude. Pitch oscillations may be as
high as "40 degrees and normally reverse direction
approximately every 3 seconds. Roll reversals up to"90 degrees from wings level may occur and the yaw
rate tends to cyclically reverse back and forth andmay be as high as 40 degrees per second. A slow net
heading change, usually to the left, may occur.
Inverted deep stalls may be either stable or highly
oscillatory in pitch, depending upon the CG. If there
is little or no pitch motion, the nose is slightly above
the horizon and the wings are generally level. If the
deep stall is oscillatory in pitch, the nose may oscillate
above and below the horizon by as much as "20
degrees. Yaw and roll oscillations in an inverted deep
stall are normally smaller than those during an
upright deep stall.

610

The aircraft is most likely to stabilize in a deep stall
if it does not selfrecover within two postdeparture
pitch oscillations. The likelihood of a deep stall
occurring after a departure is also dependent upon
the aircraft CG and configuration. An inverted deep
stall can occur at a more forward CG than an uprightone. The likelihood of a deep stall developing after adeparture increases as the CG moves farther aft and
if stores (especially 370gallon fuel tanks) are loaded.
SPIN
A spin is a deep stall with a significant sustained yaw
rate in one direction (greater than 30 degrees per
second). The pitch, roll, and yaw oscillations
associated with a deep stall should not be confusedwith the continuous yaw rotation associated with aspin. AOA, airspeed indications, and altitude loss are
similar to those during deep stalls. Spins can be either
upright or inverted, although inverted spins aremuch less likely to occur than upright spins.
The yaw rate limiter is effective in preventing an
upright spin with most CAT I loadings. However,following a yaw departure above 25,000 feet, aircraftwith CAT I loadings that have all the following
characteristics may spin:
D Centerline store.
D Inlet mounted pod(s).
D Lateral asymmetry greater than 300 pounds at

stations 1, 2, or 3.

Upright spins following a yaw departure can bedisorienting. The initial portion of the spin ischaracterized by highly oscillatory pitch and rollmotions and a high yaw rate (70 to 100 degrees per
second). Initially, the aircraft spins roughly aroundthe aircraft's flight path at departure. As the spin
continues, the rotation axis eventually becomes
vertical. Very noticeable forward g (eyeballs out) andsideforces are present.
Yaw rate usually decreases to near zero within 10 to
25 seconds. The g forces decrease noticeably as yaw
rate decreases, which may give the sensation thatyaw rate is lower than it actually is. Use outsidereferences to determine when the yaw rate has
stopped. A recovery may occur after yaw rate has
decreased to near zero. If recovery does not occur, theaircraft has settled into a deep stall and pitch rocking
must be used to recover the aircraft to controlled
flight. However, because pitch rocking is less effective
when a yaw rate is present, pitch rocking should not
begin until yaw rotation stops or is minimized (ifaltitude permits).

T.O. GR1F16CJ1
A spin may also occur following any departure with a
CAT III loading that has a large lateral asymmetry
from airtosurface stores. These spins may be fast,flat, and possibly unrecoverable.
An inverted spin can be caused by pilot rudder androll commands if they are not released following an
inverted departure.
The yaw rate limiter is effective in preventing an
inverted spin with all CAT I loadings. However, large
lateral asymmetries from airtosurface stores on
CAT III loadings may overpower the yaw rate limiterand cause unrecoverable inverted spins.
RECOVERIES
SelfRecovery
Recovery from most departures is automatic,
requiring only release of the controls. Once thecontrols are released, selfrecoveries usually occur
within the first two postdeparture pitch oscillations(1020 seconds). Recovery is characterized by the nose
pitching down to a steep dive angle, increasing
airspeed, and AOA and sideslip returning to thenormal range. Some postdeparture yaw and rolloscillations may be evident, particularly if thedeparture was in yaw. To prevent departure reentry,
the airspeed should be allowed to increase to 200
knots prior to dive recovery. Flight control failure
indications may be present after recovery from any
departure.
Deep Stall Recovery
The aircraft should be allowed the opportunity toselfrecover if altitude permits. Initiating pitch
rocking too soon should be avoided because it canaggravate postdeparture roll and yaw motions, which
can significantly lengthen recovery time.
If recovery is not apparent after two postdeparture
pitch oscillations (1020 seconds) or if altitude is afactor, the aircraft must be rocked out of a deep stall.
To pitch rock the aircraft, the MPO switch must be
firmly held in OVRD until recovery is complete.
Recognize any pitching motions and begin stick
cycling inphase with these motions. If no pitch
motions are apparent, an abrupt maximum commandstick input to pitch the nose away from the ground(full aft stick for an upright deep stall, full forward for
an inverted deep stall) will reverse the horizontal taildeflection and should generate a noticeable pitch
rate. The best indicator for timing stick cycles is thenose position relative to the horizon, or if no outsidereferences are available, the ADI may be useful. In an
upright deep stall, an aft stick command increases

the AOA and pitch angle. When the nose reaches its
highest point and reverses direction, a full forward
stick command reinforces the nosedown pitch rate.
One complete pitch rocking cycle takes approximately 6 seconds, during which time the aircraft descends
1000 to 1500 feet.
If the nosedown pitch rate is large enough, theupright deep stall is broken and AOA returns to thenormal range (below +25 degrees). If the nosedown
pitch rate is insufficient, the nose stops its downward
motion and either begins to rise or stabilizes.
Promptly reapply the full aft stick command when thenose reverses or after 23 seconds if reversal is notapparent. Holding full forward stick more than 23
seconds with the nose stabilized can generate a rapid
yaw rate and delay recovery. If the aircraft has notrecovered after 23 seconds or if a yaw rate develops,
reapply full aft stick and complete another inphasepitch rocking cycle.
Just prior to breaking the deep stall, the nosedown
pitch rate may decrease or even stop. Unless the nose
either definitely starts back up or stabilizes for longer
than 23 seconds, another pitchup cycle should not be
started. On the pitchdown cycle, the stick should beheld full forward (upright) or aft (inverted) while thenose hesitates as the stall breaks. Following a shorthesitation (less than 23 seconds), the nose continues
down to near vertical. There is frequently a distinct, lowmagnitude airframe shudder which occurs as the stallis breaking. This shudder is a favorable indication thatrecovery is occurring.
Once the deep stall is broken, aggressive pilot
commands are usually required to stop the pitch rate ina steep dive. If possible, find and track a feature on the
ground. Maintain firm pressure on the MPO switch
until airspeed reaches 200 knots. If a transition to an
opposite AOA deep stall does occur, reinforce thealready present pitch motion with the MPO stillengaged and recovery should be rapid. Transitions
during upright deep stall recoveries are most likely tooccur with 370gallon fuel tank configurations.
Recovery to controlled flight is recognized by a steep
pitch attitude (usually within 30 degrees of vertical),pitch rate stopping, and AOA in the normal range (-5
to +25 degrees). As airspeed increases above 200
knots, release the MPO switch, maintain neutral rolland yaw commands, and recover from the resulting
dive.
Stick commands during pitch rocking should be
abrupt and full command, and should reverse after
pitch motion reverses. Stick commands that are notabrupt and full command may not be effective. Rapid
fore and aft cycling of the stick out of phase with theaircraft motion is also not effective.
611

T.O. GR1F16CJ1
The number of pitch rocking cycles required forrecovery from an upright deep stall is dependent on
aircraft configuration. Generally, recovery occurs in
one or two pitch rocking cycles. However, configurations with a centerline store (particularly a
300gallon fuel tank) or with 370gallon fuel tanks
may require more pitch rocking cycles for recovery.
These loadings usually have more oscillatory deep
stalls, and the roll and yaw motions make it moredifficult to determine proper stick cycling. Pitch
attitude is still the best indication for proper stick
cycling.
During upright deep stalls with a centerline store,
particularly a 300gallon fuel tank, the aircraft tends
to roll and yaw right while pitching up, and roll andyaw left while pitching down. During deep stalls with
370gallon fuel tanks, the aircraft nose motion
appears triangular. This motion is characterized by aroll and yaw right while pitching up, followed by apitch down, a hesitation, and a yaw to the left. These
alternating yaw oscillations should not be confusedwith a sustained yaw rotation in one direction
indicating the aircraft is in a spin.
An inverted deep stall recovery is similar to an
upright recovery. In an inverted deep stall, the yaw
rate limiter automatically provides rudder against
the yaw rate. Pilot roll and rudder commands should
be avoided. Pilot roll and rudder commands are
inhibited when the MPO switch is in OVRD. Yaw
oscillations may be noticed but do not affect recovery.
To recover from an inverted deep stall, position the
MPO switch to OVRD and begin stick cyclinginphase with pitch motions. If no pitch motions areapparent, make the first pitch input away from theground by pushing full forward and monitor pitch
motion. One or two pitch rocking cycles are usually
sufficient to recover from an inverted deep stall.
During recovery from an inverted deep stall, atransition to an upright deep stall is likely to occur if
large pitch motions are present during inverted pitch
rocking and the MPO switch is released too early.
Early release of the MPO switch reduces thehorizontal tail authority available and delays
recovery. The MPO switch must be held in the OVRD
position until the deep stall is positively broken as
evidenced by the pitch rate stopping, AOA in normal
range (-5 to +25 degrees), and a steep pitch attitude(usually within 30 degrees of the vertical). As
airspeed increases above 200 knots, release the MPO
switch, maintain neutral roll and yaw commands,
and recover from the resulting dive.
612NOTE
The aircraft has demonstrated
a strong tendency to transition to an
upright deep stall during recovery
from an inverted deep stall. Aggressive pilot commands (full forward
stick) are required as the nose pitches
down to the vertical during recovery topreclude a transition.
GE 129Spin Recovery
To recover from a spin, yaw rate must be stopped orminimized before the aircraft can be recovered. Dueto large noseup moments caused by the inertial
properties of the aircraft and decreased horizontal
tail pitch effectiveness with sideslip, attempts topitch rock out of a spin are usually not effective. Pitch
rocking during a spin is also likely to aggravate rolland yaw oscillations, which make recovery moredifficult.
The large forward g (eyeballs out) and sideforces
present during the initial portion of an upright spindecrease noticeably as yaw rate decreases, which may
give the sensation that yaw rate is lower than it
actually is. Use outside references to determine when
the yaw rotation has stopped. Pitch rocking should
not begin until the yaw rotation stops or is minimized.
Pitch rocking with a steady yaw rate greater than 30
degrees per second may prevent recovery.
Waiting for yaw rotation to stop or minimize may
require from 20 to 30 seconds. When the yaw rotation
stops or is minimized, the aircraft will either
selfrecover or will settle into an upright deep stall. If
recovery is not apparent after yaw rate has stopped oris minimized, perform the appropriate recovery
procedures described in the Deep Stall Recovery
section.
Pitch, roll, and yaw oscillations associated with adeep stall should not be confused with the continuous
yaw rotation associated with a spin. Once thecontinuous yaw rotation of a spin has been arrested
and pitch rocking has begun, pitch rocking should
continue until a recovery is achieved.
Upright spins with CAT III loadings that have large
lateral asymmetries from airtosurface stores may
be fast, flat, and possibly unrecoverable. There is an
option to jettison an asymmetric store as a last ditch
effort in case of a fast, flat spin. However,aircraftstore collision may occur.
Inverted spins with CAT I loadings are effectively
prevented by the yaw rate limiter. The yaw rate limiter
automatically provides rudder against the yaw rate.
Pilot roll and rudder commands should be avoided.

T.O. GR1F16CJ1
If recovery is not apparent after yaw rate has stopped
or is minimized in an inverted spin, the aircraft has
settled into an inverted deep stall. Perform theinverted deep stall recovery procedures.
ENGINE OPERATION DURING DEPARTURES/OUTOFCONTROLPW 229 Departures at high altitude may result in anengine stall. If in AB during an outofcontrol
situation, retard the throttle to MIL. If at MIL or
below, do not move the throttle. Do not advance thethrottle until beginning the dive recovery.
GE 129 Departures at high altitude may result in anengine stall. Prolonged negative g flight at a highengine thrust level may result in an engine bearing
failure. Retard the throttle to IDLE after adeparture. Do not advance the throttle until
beginning the dive recovery.

If the engine stalls during a departure oroutofcontrol situation, refer to ENGINE STALLS,
Section III, after recovery is complete. The engine
should be left running during an outofcontrol
situation to insure adequate hydraulic pressure toflight control surfaces for recovery. Recover from theoutofcontrol condition; then concentrate on theengine. If the engine does not recover, it must be shut
down and restarted.

DEGRADED FLIGHT CONTROLSFLCS DBU
Flight characteristics for cruise and landing are notsignificantly affected by operation in DBU.
Although airspeed should be maintained below 500
knots/0.9 mach while in DBU, transitions to DBU at
higher airspeeds do not produce adverse handlingcharacteristics. Minimal pitch transients occur iftransition to DBU occurs at 1g. At higher g levels,transition to DBU is accompanied by an initial
reduction in g. If the aft stick input is continued, glevel again increases, but may be less than thatwhich was available prior to the transition. The
STORES CONFIG switch is inoperative in DBU andAOA/g force available is similar to that in CAT III.
If LG are lowered below 200 knots, a mild noseuptransient of approximately 2 degrees occurs. A
similar nosedown movement occurs if LG is raised
below 200 knots.
If an automatic transition to DBU occurs when in
auto TF, the FLCS initiates a 3g incremental no rollto wings level fly up. This fly up can be interrupted
using the paddle switch. All ATF indications areturned off (active light, TF FAIL warning light, ATF
NOT ENGAGED caution light, and autopilot PITCH
switch). HUD TF symbology is also removed.
LEADING EDGE FLAPS LOCKED (SYMMETRIC)
Flight characteristics for landing and low AOA
maneuvering are not significantly affected bylocked LEF's. At high airspeeds, LEF's locked down
cause increased buffet. At high AOA, LEF's locked
up reduce stability, increase departure susceptibil
ity significantly, and cause increased buffet. Above
1618 degrees AOA, an abrupt yaw departure may
occur, producing an uncommanded roll with little orno forewarning. Do not exceed 12 degrees AOA with
the LEF's inoperative. Locked down LEF's signifi
cantly reduce cruise range. During landing,floating may also be noticeable if LEF's are locked
at or near full down. The aircraft may float, sinkrate may decrease, and a slight forward stickpressure may be needed to fly through the groundeffect.
STANDBY GAINSWhen operating on standby gains, aircraft response
is normal at low AOA. Because the LEF's are at zero
degrees (LG handle in UP and ALT FLAPS switch in
NORM) with a dual air data failure, buffet anddeparture susceptibility will be increased at higher
AOA (above 18 degrees). At flight conditions
higher/slower than the fixed gain conditions,aircraft response is more sluggish requiring larger
control commands for a given response. Landing theaircraft should present no special problems.
ONE HYDRAULIC SYSTEM
Flight characteristics with one hydraulic systemshould be normal unless extremely large, rapid
control surface deflections are commanded. Under
these conditions, the hydraulic flow rate from theone system may be inadequate which slows down
control surface movement rates and possibly causes
sluggish aircraft response.
SPEEDBRAKES
Speedbrakes may stick fully open or open
asymmetrically. If a yawing moment is noted when
the speedbrakes are opened, close the speedbrakes.
If the speedbrakes fail to close, a significant
increase in drag results. Fully opened speedbrakes
significantly reduce cruise range.
613

T.O. GR1F16CJ1AIRCRAFT DAMAGE

LEADING EDGE FLAPS

Procedures for recovery of an aircraft with damage
will depend on the type and extent of the damage.
The following paragraphs provide general information based on analysis of past mishaps.

Damage to the LEF's may result in locked LEF's. If
the LEF's have not automatically locked, they should
be manually locked. Refer to Leading Edge Flaps
Locked (Symmetric), this section for a discussion offlight characteristics.

HORIZONTAL TAIL
Loss or partial loss of one horizontal tail surface will
not result in an uncontrollable aircraft except for
certain combinations of flight conditions (machgreater than 0.80 and altitude less than 15,000 feet),aircraft loading (heavy CAT III), and CG at or near theaft limit. Avoid abrupt maneuvering and maintainairspeed between 200 and 300 knots until landing
approach. Place the STORES CONFIG switch to CAT
III if in CAT I. Some roll stick or roll trim may berequired to maintain wings level. Use a maximum of
11 degrees AOA during final approach. Aftertouchdown, lower nose to runway as soon as practical.
FLAPERON
Separation of a flaperon from the wing (flaperon still
attached to the ISA in the fuselage) causes theoutboard end of the flaperon surface to rotate
upwards, towards the fuselage. This may be
accompanied by a roll transient. Adequate roll stickauthority should be available to counter the effect andmaintain control. With a flaperon surface separated
from one wing, landing should be made without
extending the TEF on the other wing (use alternate
landing gear extension, leave LG handle UP, andselect brakes CHAN 2). Hydraulic system pressures
should also be monitored. Upward rotation of aseparatedfromthewing flaperon surface has caused
ISA movement that resulted in a leakage failure of
hydraulic system A.
RUDDER AND VENTRAL FINS
Loss or partial loss of the rudder or a ventral fin will
not result in controllability problems unless theaircraft is above 1.5 mach. If loss or partial loss of therudder or a ventral fin occurs, avoid abrupt
maneuvering and reduce speed to subsonic, if
supersonic. Place the STORES CONFIG switch to
CAT III if in CAT I. During landing with loss or partial
loss of the rudder, lower the nose to the runway assoon as practical so that NWS can be used as requiredfor directional control.

614

WING
Loss of a portion of the wing produces a rolling motion
towards the damaged wing. The capability to stop thisroll depends on airspeed, g, and the amount of wing
surface lost. Higher airspeeds (above 250 knots) andlow g (less than 2) are essential for maximizing theamount of roll control authority available. If theaircraft is in a dive when a portion of the wing surface
is lost, apply roll stick force to stabilize the aircraft inroll before applying g to recover from the dive(altitude permitting). Use the lowest g level practicalto recover from the dive. As g is increased, additionalroll stick force is required to maintain wings level.
Depending on the amount of wing surface lost, theminimum speed to maintain adequate roll control
could be well above 200 knots.
RADOME
Loss of the radome results in loss of two air data
sources (pitot probe) and two AOA sources (AOA
transmitters); however, loss of the radome does notnecessarily cause the aircraft to be uncontrollable.
FOD to the engine and other damage to the aircraft
may also occur as the radome departs.
Loss of two air data sources will affect handling
qualities. If the two airspeed sources from the pitot
probe go to erroneously low values at the same time,
the remaining good source from the fuselage air dataprobe will be considered failed (i.e. a single failure).
In this case the FLCS will not switch to standby
gains. Gain scheduling will be based on theerroneously low airspeed; thus, the aircraft will
become increasingly pitch sensitive as airspeed isincreased. If the two airspeed sources are lost in amanner that the FLCS recognizes as two failures,
then the FLCS reverts to standby gains. In standby
gains, the opposite problem occurs. With the landing
gear up at slow speeds, the aircraft will be sluggish
compared to handling qualities prior to the radomeloss. It may seem as though the aircraft isn't
responding to stick inputs if sufficient response time
is not allowed.
Loss of two AOA sources may not be detected as a dual
failure. The end result to the FLCS is an AOA in therange of 11-13.6 degrees (i.e. no AOA limiter).

T.O. GR1F16CJ1
If the radome is lost, attempt to attain 1 g level flightat 275-300 knots and place the landing gear handle
down. Airspeed and AOA indications won't be
accurate. Use a chase aircraft to help establish speed
and keep AOA below 12 degrees. Assess the situation
by determining what failures are being annunciatedand perform a controllability check.

SAMPLE PROBLEM:

DIVE RECOVERY
Refer to figure 61. Dive recovery capability is given
as altitude lost during pullout and is a function of
pullout load factor, dive angle, true airspeed, andFLCS limiting. Plots to convert indicated airspeed
or mach number into true airspeed are provided on
the chart. Dive recovery during constant load factor
pullout may be on the AOA limiter prior to recovery,
under certain initial conditions, as airspeed isreduced. Dive recovery capability at constant loadfactor is nearly independent of store drag. The diverecovery chart is applicable to GW's between
20,000 30,000 pounds. The dive recovery chart
becomes increasingly conservative for GW's lessthan 25,000 pounds and decreasingly conservative
for GW's greater than 25,000 pounds. Increase
altitude lost during full aft stick pullout by 4
percent for each 1000 pounds in excess of 25,000
pounds GW if initial dive angle is w45 degrees andinitial airspeed is less than 500 knots.

S

GW= 25,000 pounds

A.

KIAS

= 300

B.

Initial altitude= 15,000 feetC.

Temperature

= 0°C

D.

KTAS

= 381

E.

Mach number= 0.59F.

Dive angle= 60 degrees

G.

Category/KIAS= CAT I/300

H.

Altitude lost during
limiter pullout= 2130 feet

Altitude lost during
limiter pullout= 2130 feet

For a constant g pullout, use the greater of constantg or limiter pullout altitude lost.
NOTE
The dive recovery chart is based on an
idle thrust, wings level, speedbrakes
fully open recovery. However, if airspeed is below 350 knots, altitude loss
is minimized by selecting/maintaining
MIL/AB thrust and closing speedbrakes. If airspeed is 350 knots orabove, selecting/maintaining idle
thrust and opening speedbrakes minimize altitude loss. In either case, best
dive recovery performance is obtainedby making an ADI referenced wings
level pull.

I.

J.

Maximum pullout loadfactor available= 5.8gK.

Pullout load factor

= 3g

L.

Altitude lost during
3g pullout= 4150 feet615

T.O. GR1F16CJ1Dive Recovery

13

DATA BASIS ESTIMATED

12

CONDITIONS:IDLE THRUSTNO DELAY BEFORE PULLUP INITIATIONWINGS LEVELMAXIMUM G ONSET RATEFULL SPEEDBRAKES

11
10

KNOTS
INDICATED AIRSPEED

2. Increase altitude lost during limiter pullout by
4 percent for each 1000 pounds in excess of
25,000 pounds GW if initial dive angle is 45
degrees and initial KIAS is less than 500.

6

800

5

7

L
4

K

700

7 3

600

I
6 2

J

500

5 1

400
300

A
B

4 0

LIMITER PULLOUTINITIAL KIAS

200

CAT I

250
300
400+

CAT III

3

200
250
300

H
2

G

1

400+

0

KNOTS

900

TRUE AIRSPEED

ALTITUDE LOST DURING PULLOUT

8

1. Applicable for CAT I or CAT III limiter.

1000 FEET9NOTES:

800
700
600
500
400

D

C
F

300
200
0.2

E
0.4

0.6 0.8 1.0
MACH NUMBER

1.2

1.4
1F-16X-1-0021X

Figure 61.

616

T.O. GR1F16CJ1

SECTION VII
ADVERSE WEATHER OPERATIONTABLE OF CONTENTS
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Instrument Flight Procedures . . . . . . . . . . . . . .
Holding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Penetration . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Instrument Pattern/Approaches . . . . . . . . . .
Missed Approach . . . . . . . . . . . . . . . . . . . . . . . .
Turbulence and Thunderstorms . . . . . . . . . . . . .
Cold Weather Operation . . . . . . . . . . . . . . . . . . .
Before Entering Cockpit . . . . . . . . . . . . . . . . .
Before Starting Engine . . . . . . . . . . . . . . . . . .
Starting Engine . . . . . . . . . . . . . . . . . . . . . . . . .
After Engine Start . . . . . . . . . . . . . . . . . . . . . .
Taxi . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Takeoff . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
In Flight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Landing in Icy or Wet Conditions . . . . . . . . . . .
Hot Weather and Desert Ground
Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Volcanic Ash Operation . . . . . . . . . . . . . . . . . . . .
Ground Operations . . . . . . . . . . . . . . . . . . . . . .
InFlight Operations . . . . . . . . . . . . . . . . . . . .

FThe HUD should not be used as the sole

71
71
72
72
72
72
72
75
75
75
75
75
76
76
76
76
77
77
77
78

INTRODUCTION
This section contains information and proceduresthat affect operation of the aircraft in adverse
weather and climatic conditions and which differ
from the normal procedures in Section II.

INSTRUMENT FLIGHT PROCEDURESPX II The HUD may be used as a reference for
instrument flight. PX III The HUD may be used as a
primary reference for instrument flight. Frequentcrosschecks with other instruments will be per
formed to maintain proper flight orientation anddetect failures that are not directly communicated tothe pilot.

F PX III The EGI is not certified toprovide instrument approach flightpath guidance.

reference for instrument flight due tothe lack of adequate failure warning
but should be crosschecked with theprimary/basic instruments.

F PX III The displays generated for the

JHMCS helmet are not approved foruse as a reference during instrument
meteorological conditions (IMC) or forcourse guidance during landing.

F A delayed selection of ILS/TCN or

ILS/NAV until the aircraft is nearly on
the ILS glide slope may cause the flight
director circle to be positioned incorrectly(full up or full down) for up to 90 seconds.
If the flight director circle is positioned
incorrectly when an ILS mode isselected, move the INSTR MODE knob
to TCN or NAV, then back to the desired
ILS mode. This action enables the flight
director to operate properly.

FIt is possible for the displayed ADI and/or HUD attitude to be in error with no
ADI OFF or AUX warning flags in view
and without an INS or HUD PFL.
Displayed HSI and/or HUD headings
may also be in error with no HSI OFF
or ADI AUX warning flags in view andwithout an INS or HUD PFL. Momentary warning flags may indicate
impending failure. To detect thesefailures and maintain proper flightorientation, basic and backup instru
ments shall be crosschecked.

F PX III It is possible for the MMC toposition the ILS glideslope bar full down
without being dashed even though the
ILS glideslope signal may be valid. Caremust be taken during precision ap
proaches to cross check ILS informationon the ADI if the HUD glideslope bar
drives full down and is not dashed.
71

T.O. GR1F16CJ1PENETRATIONF PX II With certain failures of the INS

with ILS selected, a fixed aircraft
reference symbol is displayed at zero
degrees azimuth, 11 degrees below theboresight symbol. This symbol is for ILS
deviation reference in conjunction withthe horizontal and vertical deviation
bars. The aircraft reference symbol
should not be used for attitude reference.

FAvoid touching the canopy transparency, canopy frame or placing arms on the
canopy body positioning handles during IMC. Contact may produce severe
shock as a result of static discharge.
NOTE

FWhen rotating the HSI course settingfrom an eightunit digit to a nineunit
digit (e.g., 018 to 019), the tens digit
may rotate prematurely causing a
10degree reading error (e.g., 029).
Crosscheck counter setting withcourse arrow reading to insure propercourse setting.

FElectricaltransients (particularly
during the EPU check) may cause the
INS to revert to use of the last
manually entered magnetic variation.
Under these conditions, the accuracyof heading information displayed on
the HSI and HUD depends upon the
manual magnetic variation value
entered. If magnetic heading error issuspected, confirm that automagnetic
variation is in use or that the correct
local magnetic variation is entered.

F PX III The EGI does not use navigational

aid assigned variation to compute course
display for TACAN, VORTAC, VOR, or
VORDME points retrieved from thedatabase. Rather, the magnetic variation
at the aircraft present position is used forcourse computation and display.

Penetrations are normally flown at 300 knots, speedbrakes as required, and throttle at IDLE.
INSTRUMENT PATTERN/APPROACHES
Refer to figure 71 for a typical TACAN holding,penetration, and approach. Refer to figure 72 for atypical GCA. Instrument patterns are normallyflown at 200250 knots clean. Approaching final,
lower the LG, slow the aircraft, and fly final
approach at 13 degrees AOA maximum.
MISSED APPROACH
Advance throttle as required, close speedbrakes, andretract LG after a positive climb is established. Adjust throttle to maintain between 200250 knots.
Pitch transients resulting from LG and TEF changes
are mild and require minimum control compensation.

TURBULENCE AND THUNDERSTORMS
Avoid flight in turbulent air, hailstorms, andthunderstorms. There is a high probability ofdamage to airframe and components from impact
ice, hail, and lightning. Thunderstorm penetration
airspeed is 300 knots or optimum cruise airspeed,whichever is lower. At high airspeeds, personal
discomfort and structural stress are greater. At
slower airspeeds, controllability is reduced and inlet
airflow distortion due to turbulence may cause
compressor stall and/or engine stagnation.
The GM mode of the radar can be used as an aid innavigation between or around storm cells. Refer toT.O. GR1F16CJ3411 for expanded information. If
entry into adverse weather cannot be avoided, thefollowing procedures should be used:
1.

PROBE HEAT switch - Check PROBE HEAT.

2.

FLOOD CONSOLES knob - HIGH INT.

3.

ANTI ICE switch - ON.

4.

Airspeed - 300 knots or optimum cruise,
whichever is lower.
NOTE

HOLDING
Holding airspeed is 200250 knots (maximum
endurance airspeed recommended).
72

Severe turbulence causes variations inairspeed and altitude. Do not change
throttle setting except for extreme
airspeed variations.

T.O. GR1F16CJ1TACAN Holding, Penetration, and Approach(Typical)

HOLDING (ALL ALTITUDES)MISSED APPROACH

AIRSPEED

200-250 KNOTS

THROTTLE
AS REQUIREDSPEEDBRAKES
CLOSE
LG
UP
AIRSPEEDACCELERATETO 200-250 KNOTS

INITIALAPPROACH FIXFINALAPPROACH FIXTACAN APPROACHPENETRATION DESCENTLG
DN
THROTTLE
AS REQUIREDAOA
13 DEGREES (MAX)SPEEDBRAKESOPEN

AIRSPEED
300 KNOTS
SPEEDBRAKESAS REQUIREDTHROTTLE
IDLE
ALTIMETER
SET

NOTE:
A typical straight-in penetration and approach require approximately 400 pounds of fuel for a drag index of 0 and
600 pounds of fuel for a drag index of 200.

GR1F-16CJ-1-0132X37

Figure 71.

73

T.O. GR1F16CJ1

GCA (Typical)

BASE
LG
AS REQUIREDSPEEDBRAKESASREQUIRED
AOA
13 DEGREES(MAX) (LGDN)

ENTRY/DOWNWINDAIRSPEED
200-250
KNOTS (LG
UP)

FINAL
LG
CHECK DN
SPEEDBRAKES
OPEN
AOA
13 DEGREES(MAX)MISSED APPROACH
THROTTLE
AS REQUIREDSPEEDBRAKES
CLOSE
LG
UP
AIRSPEEDACCELERATETO 200-250 KNOTS

NOTE:
A typical GCA pattern requires approximately 400 pounds of fuel for a drag index of 0 and 700 pounds of fuel fora drag index of 200.

GR1F-16CJ-1-0133X37

Figure 72.

74

T.O. GR1F16CJ1NOTEFAn extremely loud screeching noise

may be heard in the headset while
flying in cirrus clouds or in the vicinityof thunderstorms. The noise may beeliminated by turning the UHF or VHF
radio off, by turning the volume(s)down, or by changing UHF antenna
positions.

FWhen flying in heavy rain, water tends

to be aerodynamically held on theforward portion of the canopy. At
higher airspeeds, this condition may
obscure visibility as much as 30
degrees back on each side of the
canopy. On final approach, the water isgenerally confined to the position of
the canopy immediately in front of the
HUD. It may be necessary to look out
the sides of the canopy to acquire therunway and to flare and land theaircraft.

COLD WEATHER OPERATION

Engine operation under the following
conditions may result in engine dam
age due to icing:
D Ambienttemperatures between
20_F (-7_C) and 45_F (7_C) with
precipitation (rain, fog, sleet, orsnow).

D Dewpoint

within 9_F (5_C) of
ambient temperatures between
25_F (-4_C) and 45_F (7_C).

D Ambient temperature below 45_F

(7_C) with standing water or a
mixture of water with ice or snow
within the immediate proximity ofthe engine inlet.

BEFORE ENTERING COCKPIT
All accumulated ice and snow must be removed fromthe aircraft before flight is attempted. Insure thatwater does not accumulate on control surfaces orother critical areas where refreezing may cause
damage or binding.

Do not permit ice to be chipped or
scraped away.
BEFORE STARTING ENGINE
Extreme cold temperature may require cockpit
preheating to ease operation of rotarytype switches.
D The canopy may not latch on battery power alone.
Start the engine with the canopy closed as much as
possible.
If there is visible moisture and ambient temperature
is 45_F (7_C) or less, place the ANTI ICE switch to
ON. This reduces ice buildup on the engine front
face, eliminates ice on the heated inlet strut, andreduces the possibility of ice ingestion.
STARTING ENGINE
If the aircraft is serviced with MILH5606 hydraulicfluid and the aircraft has cold soaked for more than
1 hour at temperatures below -40_F (-40_C), do notstart the JFS until ambient temperature increases
to above -40_F (-40_C) for at least 2 hours or until
the engine bay is preheated. For temperatures above-40_F (-40_C), refer to JET FUEL STARTER
LIMITS, Section V.
If the aircraft is serviced with MILH83282
hydraulic fluid and the aircraft has cold soaked formore than 1 hour at temperatures below -20_F(-29_C), do not start the JFS until ambient
temperature increases above -20_F (-29_C) for atleast 2 hours or until the engine bay is preheated. For
temperatures above -20_F (-29_C), refer to JET
FUEL STARTER LIMITS, Section V.
During cold start, oil pressure may be 100 psi for up
to PW 229 1 minute, GE 129 2 minutes.
AFTER ENGINE STARTEPU fuel quantity can indicate as low as 90 percent
at temperatures below 40_F (4_C).
For rapid cockpit warming, position the TEMP knob
to the desired MAN WARM range. Position the
RADAR to off and the DEFOG lever as required toclear fogging. After the cockpit reaches a comfort
able temperature, select a setting within the AUTO
range. If the engine was started with the canopy
unlatched, wait approximately 10 minutes to warm
the canopy before fully closing it.
If probe icing is evident or suspected, turn the
PROBE HEAT switch to PROBE HEAT at least 2
minutes prior to accomplishing the FLCS BIT.
75

T.O. GR1F16CJ1TAKEOFF

If probe heat is on or has been on, heatin probes may be sufficient to cause
injury if touched.

If the aircraft has cold soaked attemperatures below -20_F (-29_C),
repeated brake applications (2530)
may be required before the brakes
work effectively.
TAXI
To avoid brake icing, do not taxi in deep water, slush,or deep snow. When taxiing on ice or hard packed
snow, NWS may not be completely effective. Use a
combination of NWS and differential braking tomaintain directional control. Taxi at a safe speed
considering surface condition, GW, slope, and thrust.
If the aircraft has cold soaked at temperatures below-20_F (-29_C), the NWS may initially be sluggish,
but controllable.

Probe internal icing must be suspectedanytime the aircraft has been exposed
to near or below freezing conditions on
the ground. Internal icing may bedifficult to see and may remain present
even when current conditions do notappear conducive to ice formation.
Turn probe heat on at least 2 minutes
prior to takeoff anytime icing of probes
is possible.

If unable to control taxi speed or
direction, immediately shut down theengine.
NOTE
After cold soaking at temperatures
below 0_F (-18_C), be alert for flat
MLG struts.
76

If the aircraft has cold soaked at temperatures below-20_F (-29_C), LG retraction times may besignificantly increased. In addition, the nose geardoor may fail to close. If the nose gear door can bevisually confirmed as the only LG component that has
failed to retract/close, then up to two extend/retract
cycles can be made in an attempt to achieve a normal
LG up condition. Observe LG extended or in transit
limitations, Section V.
IN FLIGHT
Flight in areas of icing should be avoided wheneverpossible. If icing conditions are anticipated or cannot
be avoided, turn ANTI ICE switch to ON and PROBEHEAT switch to PROBE HEAT. Frequently checkthe aircraft leading edges for indication of ice
buildup. Make all throttle movements slower thannormal when in potential icing conditions to reduce
possibility of engine stalls and/or stagnation.
Consider diverting to an alternate field if required to
avoid icing conditions.

LANDING IN ICY OR WET CONDI
TIONS
Icy or wet runway conditions may pose severe
problems in directional control and braking
effectiveness due to hydroplaning. Althoughpossible, total hydroplaning is not expected below
130 knots groundspeed. Partial hydroplaning canoccur to varying degrees below 130 knots. Oncehydroplaning occurs, it can continue to speeds well
below the onset speed. Wheel spinup must occur topermit normal antiskid braking. Hydroplaning canprevent wheel spinup and can occur on runways
which only appear damp if heavy braking is applied
at high speeds. Hydroplaning tendency increases
with water depth and with smooth runway surfaces
such as rubber deposits or paint stripes.
LESS b2t Approach and touchdown are the same as
for a short field landing on a dry runway.
Immediately after touchdown, make a deliberate
effort to be sure brakes are not applied while using
the rudder. Deploy the drag chute (if required)
immediately after touchdown. Use twopoint
aerodynamic braking until approximately 100
knots; then fly the nosewheel to the runway.
Maximum effective twopoint aerodynamic braking
is achieved at 13 degrees AOA. After the nosewheel
is on the runway, open the speedbrakes fully andmaintain full aft stick for maximum threepoint
aerodynamic braking and wheel braking effective
ness. Test for braking effectiveness before using full

T.O. GR1F16CJ1
continuous braking by momentarily depressing
pedals, fully releasing pedals for at least onehalf
second, and then depressing pedals again. This
technique gives the wheels a better opportunity tospin up if hydroplaning conditions exist. If braking
effectiveness is not felt, continue to pump brakes as
speed decreases, making sure pedals are momentarily fully released between applications. Usecontinuous braking after braking effectiveness isfelt. As speed decreases, the antiskid system will
increase deceleration accordingly.
LESS b2t When stopping distance is critical, wheel
braking should be initiated when below 100 knots.
Wheel braking effectiveness at high speeds is very
low compared to twopoint aerodynamic braking
effectiveness. Low deceleration at high speed may be
mistakenly interpreted as a brake or antiskid
failure. If braking effectiveness or anti skid cycling
is not perceived, release brakes momentarily andthen reapply brakes. When the wheel brakes become
effective, the nose will automatically lower. After the
nosewheel is on the runway, maintain full aft stick,open the speedbrakes fully, and use continuous
maximum braking. Do not hesitate to lower thehook, if required.
LESS b2t If crosswinds preclude maintainingtwopoint aerodynamic braking, test for braking
effectiveness as previously discussed before using
full continuous braking. Continue to pump thebrakes until braking effectiveness is felt and speed
is below 100 knots.
b2t Approach and touchdown are the same as for ashort field landing on a dry runway. Immediately
after touchdown, make a deliberate effort to be surebrakes are not applied while using the rudder. Use
twopoint aerodynamic braking until approximately
100 knots; then fly the nosewheel to the runway.
Maximum effective twopoint aerodynamic braking isachieved at 13 degrees AOA. After the nosewheel is on
the runway, open the speedbrakes fully and maintainfull aft stick for maximum threepoint aerodynamic
braking and wheel braking effectiveness. When
wheel brakes are used, they should be continuously
applied.
b2t When stopping distance is critical, continuous
maximum wheel braking should be initiated in thetwopoint attitude. Wheel braking effectiveness athigh speeds is very low compared to twopoint
aerodynamic braking effectiveness. Low deceleration
at high speed may be mistakenly interpreted as abrake or antiskid failure. When the wheel brakes
become effective, the nose will automatically lower.
Do not hesitate to lower the hook, if required.

FLESS b2t Continuous wheel braking

above 100 knots is not recommended.
Hydroplaning may prevent spinup ofboth MLG wheels and wheel brakes
may become operative without anti
skid protection. Locked wheels andsubsequent blown tires can occur.

FRubber deposits on the last 2000 feet ofa wet runway make directional control
a difficult problem even at very low
speeds. Braking should be started in
sufficient time to avoid excessive
braking on the last portion of the
runway.

HOT WEATHER AND DESERT GROUNDOPERATION
Hot weather and desert ground operation requiresthat added precautions be taken against damage
from dust, sand, and high temperatures. Particularattention should be given to those components andsystems (engine, fuel, oil, hydraulic, pitotstatic,etc.) which are susceptible to contamination,
malfunction, or damage from sand and dust. Inspect
the pistons on the LG and have them cleaned as
required. Check the engine inlet duct for sandaccumulation. During conditions of blowing sandand dust, the canopy should be closed and sealed andall protective covers installed when the aircraft is
not in use.
In hot, humid conditions, fogging of the exterior
canopy surface after flight may reduce visibility tothe point where the canopy must be opened prior totaxiing. Stow all equipment prior to opening the
canopy.

VOLCANIC ASH OPERATIONGROUND OPERATIONS
Modified ground operations on an airfield which has
experienced volcanic ash fallout are required evenafter cleanup is complete.

77

T.O. GR1F16CJ1
For preflight, carefully inspect the following areasfor volcanic dust:
D ECS ram inlet ducts.
D Engine inlet.
D LG strut chrome surface.
D Pitot tube, air data probes, and AOA probes.

Some indications of volcanic ash cloud penetrationare:D Acrid odor.
D Canopy opaquing.
D Engine exhaust torching.

D Static ports.

After engine start, keep ground operation time andthrust to a minimum and run airconditioning at fullcold setting, if practical. Do not use antiice unlessrequired.
Taxi at a safe speed considering surface conditions,
GW, slope, and thrust. As volcanic ash, especiallywhen wet, reduces RCR, perform a rolling takeoff, if
possible.
Consider increasing the interval between takeoffs toallow clouds of suspended volcanic ash to clear.
After landing, consider shutting down the engine
and have aircraft towed to the parking area.
Use RCR of 18 for dry volcanic ash and 10 for wet
volcanic ash if actual RCR value is unknown.
INFLIGHT OPERATIONSFlight in a volcanic ash environment is extremely
hazardous. Airborne radar does not detect ash
clouds which visually are easily mistaken for normal
clouds.

78

D Engine surges/malfunctions.
D Erroneous airspeed indications/fluctuations.
D FTIT rise.
D Saint Elmo's fire.
D Volcanic ash entering cockpit.

Upon detecting volcanic ash cloud penetration,
reduce thrust and maintain minimum thrust
required to sustain flight, exit the volcanic ash
environment, and land as soon as practical. Do not
use antiice unless required.
PW 229 Engine operation above 80 percent rpm while
ingesting volcanic ash may cause buildup of melted
ash on turbine hardware and possible engine stalls.
If possible, keep rpm at 80 percent or less until clearof volcanic ash.

T.O. GR1F16CJ1

SECTION VIII
AIR REFUELING PROCEDURESTABLE OF CONTENTS

Page

Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Control of Tanker/Receiver Forces . . . . . . .
Wingman Receiver Responsibilities . . . . . .
Airspeeds and Altitudes . . . . . . . . . . . . . . . .
Fuel Reserve Requirements . . . . . . . . . . . . .
Weather . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Refueling Transfer Rate . . . . . . . . . . . . . . . .
Communications . . . . . . . . . . . . . . . . . . . . . . .
Oral Communications . . . . . . . . . . . . . . . .
Visual Signals . . . . . . . . . . . . . . . . . . . . . . .
Hot Armament Procedures . . . . . . . . . . .
Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Receiver Director Lights KC135 . . . . . .
Receiver Director Lights
KC10/KDC10 . . . . . . . . . . . . . . . . . . . .
Navigation and Position Reporting . . . . . .
Post Air Refueling . . . . . . . . . . . . . . . . . . .

82
82
82
82
82
82
82
82
82
83
86
86
86
811
811
811
812

Flying Safety . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Boom Envelope Limits . . . . . . . . . . . . . . . . . .

812
812

Air Refueling Procedures . . . . . . . . . . . . . . . . . .
Departure/Enroute Procedures . . . . . . . . . .
Cell Leader Responsibility . . . . . . . . . . . .
Buddy Departure . . . . . . . . . . . . . . . . . . . .
Taxi . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
LineUp . . . . . . . . . . . . . . . . . . . . . . . . . .
Takeoff . . . . . . . . . . . . . . . . . . . . . . . . . . .
Aborts During Takeoff . . . . . . . . . . . . .
Element JoinUp and Climb . . . . . . . .
Element Cruise Formation . . . . . . . . . . .
Force Extension Procedures . . . . . . . . . .
VMC Procedures . . . . . . . . . . . . . . . . . .
IMC or Night Procedures . . . . . . . . . .
Post Refueling Procedures . . . . . . . . .
Lost Wingman Procedures . . . . . . . . . . . .
Abort Procedures . . . . . . . . . . . . . . . . . . . .
Rendezvous Procedures . . . . . . . . . . . . . . . . .
Altimeter Settings . . . . . . . . . . . . . . . . . . .
Rendezvous Altitude Block . . . . . . . . . . .
Track . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Tanker Rendezvous Equipment . . . . . . .
Receiver Formation During
Rendezvous . . . . . . . . . . . . . . . . . . . . . . .
Early Arrival . . . . . . . . . . . . . . . . . . . . . . . .
Tanker Identification . . . . . . . . . . . . . . . . .

817
817
817
817
817
817
817
817
817
818
818
819
819
819
819
820
820
820
820
821
821
821
822
822

Point Parallel Rendezvous . . . . . . . . . . . .
Point Parallel Rendezvous With
Tanker Escort . . . . . . . . . . . . . . . . . . . . .
Receiver TurnOn Rendezvous . . . . . . . .
Rendezvous Overrun . . . . . . . . . . . . . . . . .
Enroute Rendezvous . . . . . . . . . . . . . . . . .
Departure and Climb . . . . . . . . . . . . . .
Enroute Rendezvous Procedures . . . .
Alternate Rendezvous Procedures . . . . .
Emission Option 3 . . . . . . . . . . . . . . . . . . .
Missed Rendezvous Procedures . . . . . . .
Special Air Refueling Procedures . . . . . . . .
Refueling Sequence . . . . . . . . . . . . . . . . . .
Fingertip Formation . . . . . . . . . . . . . . .
Echelon Formation (VMC Only) . . . .
Fuel Management . . . . . . . . . . . . . . . . . . .
Precontact . . . . . . . . . . . . . . . . . . . . . . . . . .
Boom and Receptacle Procedures . . . . . .
Disconnect KC135 . . . . . . . . . . . . . . . .
Disconnect KC10/KDC10 . . . . . . . . .
Quick Flow Air Refueling
Procedures . . . . . . . . . . . . . . . . . . . . . . . .
Toboggan . . . . . . . . . . . . . . . . . . . . . . . . . . .
Weather Abort Procedures . . . . . . . . . . . .
Afterburner Air Refueling . . . . . . . . . . . .
Separation/Termination Procedures . . .
Separation From a Single Tanker . . .
Cruise Cell Termination (VMC) . . . . .
Cruise Cell Termination (IMC) . . . . .
Cell Termination at Terminal
Approach Fix . . . . . . . . . . . . . . . . . . .
Element Penetration . . . . . . . . . . . . . .
Receiver Radar Rejoin Procedures(IMC) . . . . . . . . . . . . . . . . . . . . . . . . . .
Normal Air Refueling Procedures . . . . . . . .
Armament Safety Check . . . . . . . . . . . . .
Precontact . . . . . . . . . . . . . . . . . . . . . . . . . .
Contact . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Disconnect . . . . . . . . . . . . . . . . . . . . . . . . . .
Post Air Refueling . . . . . . . . . . . . . . . . . . .
Emergency Air Refueling Procedures . . . .
Breakaway Procedures . . . . . . . . . . . . . . .
System Malfunctions . . . . . . . . . . . . . . . . .
Slipway Door Will Not Open . . . . . . .
Slipway Door Will Not Close . . . . . . .
Inoperative Boom/Receptacle
Latching . . . . . . . . . . . . . . . . . . . . . .
KC10/KDC10 Boom FLCS Failure . . .
Brute Force Disconnect . . . . . . . . . . . . . .
Inadvertent Disconnect . . . . . . . . . . .
Controlled Tension Disconnect . . . . .

822
823
824
824
824
824
824
826
827
827
827
827
827
827
828
828
828
829
829
829
831
831
831
831
831
831
831
831
832
832
834
834
834
834
834
834
835
835
835
835
835
836
836
836
836
836

Change 181

T.O. GR1F16CJ1INTRODUCTION
This section contains information and procedures forF16 air refueling with KC135 and KC10/KDC10
aircraft. For basic flight crew air refueling procedures, refer to T.O. 11C1.

GENERAL
This section reflects Emission Option 2 procedures
unless noted within the text. Inflight situations andsound judgement may dictate discontinuing commu
nications procedures outlined for Emission Option 2.

altitudes may be required for abnormally high free airtemperatures. The controlling agency directing themission will be responsible for obtaining enroute andair refueling altitude clearance for training andoperational missions.
FUEL RESERVE REQUIREMENTS
For deployment operations, the last receiver in thecell will depart the penetration fix at the abort ordestination base with a minimum of 30 minutes offuel remaining, computed in accordance withcommand guidance.

Both tanker and receiver crew must be thoroughlyfamiliar with all aspects of the refueling in order toadequately plan the mission. Planners will coordi
nate and crews will be thoroughly familiar withmission requirements as prescribed in the appropri
ate command directives.

WEATHERThe air refueling operation requires precise anddetailed planning to insure success. Each receiver unit
will maintain a file of T.O. 11C1 along with thismanual.

Rendezvous and air refueling will not be attempted
when inflight visibility is deemed insufficient for safe
air refueling operations; however, the aircraft may
close to the lockon limits of the radar provided thatthe required altitude separation is maintained until
visual contact has been established. Without lockoncapability, minimum visibility for rendezvous is 1 NM.

CONTROL OF TANKER/RECEIVER FORCES
An airborne tanker force commander and alternate
commanders, as required by the mission, will be
designated for each air refueling area. During
operational missions, the tanker commander is incommand of the air refueling operation from theperiod subsequent to positive radio contact between
the tanker cell leader and the receiver leader during
rendezvous until the end of the refueling or
termination of route cell formation, as applicable. The
airborne tanker force commander will coordinate
with the receiver force commander to insure
successful mission completion.

Weather minimums are prescribed by command
guidance. Buddy departure minimums are 1500 feetand 3 NM for day and 2500 feet and 3 NM for nighttakeoffs.

REFUELING TRANSFER RATE
The air refueling transfer rate averages 2,000 pounds
per minute with two tanker A/R pumps operatingwith a KC135 and 3,000 pounds per minute with aKC10/KDC10.
COMMUNICATIONSWINGMAN RECEIVER RESPONSIBILITIESTo assist the cell leader in insuring the safety andintegrity of the flight, the wingman receiver will:
D Keep the leader in visual or electronic contact at all

times.

D Maintain briefed position at all times.
D Anticipate corrections/changes and plan accordingly.
D Monitor all aspects of formation operations andadvise the cell leader if an unsafe condition is noted.

AIRSPEEDS AND ALTITUDES
Cruise and air refueling KCAS is 310 knots at 30,000
feet unless specifically directed otherwise. Lower
82Change 1

Except during an emergency fuel situation, air refueling operations will not be
conducted when radio communication
capability is lost between tanker andreceivers. If radio communications are
lost, or unreadable between the boom
operator and receiver pilot, contacts will
not be attempted.
Emission Option 2 will be used as the normal
rendezvous and air refueling procedures. Emission
Option 2, 3, or 4 procedures do not preclude verbal
communications for safety of flight situations or toinsure mission success. Boom interphone should be
used when compatible.

T.O. GR1F16CJ1Communications procedures and plans for rendez
vous and air refueling as outlined in pertinentcommand directives will apply. Deviations must be
specifically authorized by the appropriate command
headquarters.

Tanker(s) and/or receiver(s) will make an additional
radio call confirming level at the proper rendezvous
altitude if they are not at the proper rendezvous
altitude when the 15 minute prior to the rendezvous
control time call is made.

Unless directed otherwise, communication capability
between tankers and receivers will be maintained
during all normal rendezvous and air refueling
operations. Voice transmissions, however, will be held
to an absolute minimum during rendezvous and air
refueling to be in accordance with the Emission Option
being used.

NOTE

A TalkThroughTheBoom" interphone system isavailable after contact when air refueling with all
KC135 tankers if hot mic is selected.
KC10/KDC10 aircraft are equipped with UHF, HF,
and VHF radios and a boom interphone system.
All crewmembers must be thoroughly familiar withall required oral, visual, and electronic means ofcommunications. Strict radio discipline must be
adhered to at all times. All calls will be prefaced withindividual call signs. Tankers will begin monitoring
designated frequencies and will have the Radar
Rendezvous Beacon operating at least 30 minutes
prior to the rendezvous control time. The A/A TACAN
will be tuned to the appropriate channel 15 minutes
prior to the rendezvous control time unless it isrequired for navigational purposes. Receivers willcall 15 minutes prior to the air refueling control time,advising the tanker(s) of call signs, any changes in
ETA (minutes early or late) or altitude, and hotarmament check (if required).
NOTE
F If tankers and receivers are in contact

with a common facility providing rendez
vous assistance, radio contact between
the tanker and receivers may be delayed
to accomplish the rendezvous.

FDuring enroute rendezvous, all A/Requipment operations, interplane
communications, and timing should bebased on the RZ time. For example, the
A/A TACAN should be tuned to theappropriate channel 15 minutes prior
to the RZIP unless it is required fornavigational purposes.

The tanker will advise the receiver(s) of their callsign, air refueling altitude and, if applicable, anychange in tanker timing that would affect therendezvous (in minutes early or late).

Tankers and receivers will include
altimeter setting with appropriate
altitude calls if other than 29.92 is used.
For example, RENO 01, ONE TWO
THOUSAND FEET, ALTIMETER SETTING THREE ZERO ZERO FOUR, ON
TIME." If EMCON 3 or 4, altimetersetting must be prebriefed.
For all rendezvous and air refueling operations,tankers and receivers will normally use theirindividual flight call signs unless directed otherwise
in operational plans. When assured no other counit
formation will be in range of or using the frequency,
and/or a discrete tactical frequency has been assigned
to the formation, flight call signs may be abbreviatedfor clarity and brevity; for example, RENO
FLIGHT....GO ECHELON" (acknowledge) TWO"
THREE."
Mandatory calls for the receivers are as follows:
D Initial radio call 15 minutes prior to the rendezvous

control time.

D Notify the tanker when established on the proper

rendezvous altitude, if not at the proper rendezvous
altitude at the 15 minute prior to the ARCT call.

D Precontact call (required by Flight Leader only).

Oral Communications
NOTE
FWith the exception of the breakaway

calls, crewmembers may shorten indi
vidual flight call signs using only the
number. Example: Tank 11" would be11."FNormally,

the receiver leader willproceed to the precontact position. Whenthe leader has completed refueling,
subsequent receivers will move from theobservation position as precoordinated.

The communications requirements should be estab
lished prior to the flight. For different EMCON levels,
refer to figure 81 for Emission Option Communica
tions and figure 82 for Emission Option Emitters.

Change 183

T.O. GR1F16CJ1Emission Option Communications
ITEM

EMISSION OPTIONACTION

1

2

3

4

*

**

1.

Radios Set 30 Minutes Prior to ARCT (if dual radio capable)

X

X

2.

15 Minute Call

X

X

3.

A/A TACAN Set 15 Minutes Prior to ARCT

X

X

4.

Beacon Positive Identification (if applicable)

X

5.

ADF Check (if applicable)

X

6.

Halfway Through Turn Call (Tanker)

X

7.

Mandatory Boom Operator Calls
a. Precontact Callb. Clear Receiver to Contact
c. Acknowledge Contact/Disconnect
d. Verbal Corrections
e. Advise Receiver(s) to Return to Precontact for checklist
or equipment considerations

8.

Mandatory Receiver Calls After 15 Minute Calla. Visual Contact Established/Lost to Include Overrun
b. Precontact Callc. When Contact or Disconnect is Maded. Verbally notify Boom Operator prior to Manual/
Emergency Boom Latching Procedures

X

***

X

X
X
X
X

X
X

X

X
X

X
X

9.

Post Air Refueling ReportX

10.

1 Mile Closure Call (Receiver)

X*Radio silent. Use of other emitters is authorized unless prohibited by Supported Operations Plans.
**No emissions (UHF, VHF, TACAN, FCR/Radar, AIFF/IFF, exterior lighting, etc.) unless authorized by Air
Tasking Order, Rules of Engagement, Operations Plans, Safe Passage procedures, or other mission
directives.
***Point parallel rendezvous only.
NOTE:
Variations may be indicated by, EMCON 2. Item 7a/8b COMM N/A." This would mean normal Emission Option
2 procedures, except the precontact call would be deleted.

Figure 81.

84Change 1

T.O. GR1F16CJ1Emission Option Emitters
ITEM

EQUIPMENTEMISSION OPTION
1

2

3

4

1.

FCR/RadarOn

On

As required

Off

2.

CARA/Radar AltimeterOn

On

As required

Off

3.

TACAN/DMEOn

On

As required

Off

4.

AIFF/IFFOn

On

As required

Off

5.

UHF/VHF RadioOn

On

MonitorMonitor

6.

Exterior LightingOn

On

As required

Off

NOTE:
Variations may be indicated by, EMCON 3. Item 1 Emitters Off." This would mean normal Emission Option
3 procedures, except the FCR would be off.

Figure 82.

Change 185

T.O. GR1F16CJ1Normally, boom visual signals will be used
exclusively; however, if required or requested by thereceiver, the boom operator will begin communica
tions when the receiver reaches approximately 50
feet from the contact position. Direction, if required,
will precede distance for receiver to move and will begiven until the receiver reaches the contact position.
Example: Forward 50," Up 4," Back 2." Whencontact is established, the tanker will state, (Tanker
call sign), contact."
For Emission Option 1 and 2, the boom operator will
make a precontact audio check with receiver(s) andthe receiver(s) will acknowledge. Example: Tanker
will say, 25/57." The receiver will reply, 25."
During receiver pilot demonstration of air refuelingenvelope limits, the boom operator will state boom
limit and give the boom position for the limit being
demonstrated in increments of 2.
When the tanker is required to use manual operation,
without disconnect capability, the boom operator willstate: (Receiver call sign), the following contacts will
be made in tanker manual operation. Receiver air
refueling system will remain in normal and receiverpilot must initiate all disconnects. (Tanker call sign),
ready." Receiver pilot acknowledges by stating,(Receiver call sign), ready."
Visual Signals
Refer to figure 83. Radio silent air refueling can be
conducted by use of visual signals provided thefollowing precautions and procedures are observed:
D The method, time and place of rendezvous, andamount of fuel to be transferred must be covered in
the briefing of each crew.

D The tanker will use the receiver director lights (redonly) to aid in positioning the receiver. A steady red
light indicates a large correction and a flashing red
light indicates a small correction in the direction
indicated.

D If the need for an emergency breakaway occurs

during radio silent air refueling, oral breakaway
procedures and the visual signal in figure 83 will be
used.

Hot Armament Procedures
Prior to rendezvous with the tanker for air refueling,receiver aircraft carrying forward firing ordnancewill conduct an Armament Safety Check inaccordance with NORMAL AIR REFUELING

86Change 1

PROCEDURES, this section. When radio silence ismandatory, the receiver leader will conduct a visual
challenge with each member of his flight by pointing
his index finger straight forward and thumb upward(simulating a pistol). Each member of the flight willcomplete the safety check and respond to the leader
by raising his hand and showing a circle formed by his
index finger and thumb. To reduce the possibility of
inadvertent firing, receivers will not reposition any
electrical switches while behind a tanker unless thoseswitch changes are required for air refueling
operations or aircraft control.
LIGHTING
Refer to figure 84. While approaching the precontact
or contact position, the receiver pilot can adjust
exterior lighting as required by the boom operator.
NOTE
Visual contact for night air refueling canbe aided by requesting the tanker toflash his landing lights prior to and/or
during the tanker turn.
On single KC135 tankers performing a rendezvous
both upper and lower strobes will display red.
NOTE
If the spare is used during the air
refueling, the appropriate color code will
be displayed until the receiver is in theprecontact position. To further aid in
identification, tanker position lights will
be placed on BRIGHT and FLASHING
for numbers 1, 3, and 5. Position lights
for numbers 2 and 4 will be BRIGHT and
STEADY. Position lights will be set prior
to takeoff. After the receiver has
established visual contact and has closed
to 1/2 NM in trail, tankers will turn
position lights to STEADY and DIM andturn strobe light OFF. When any aircraft
will be flying visual wing formation on
the tanker, the tanker will also turn off
the strobe light. In this case, the last(outside) receiver aircraft with eachtanker will have anticollision lights ON.
When fighter receivers reach the obser
vation position, tankers will turn under
wing, underbody, and nacelle illuminat
ing lights to DIM. Exterior lights willthen be adjusted as requested by thereceiver pilot.

T.O. GR1F16CJ1Visual Signals
INDICATION
SIGNALBOOM AIR REFUELINGPROBE & DROGUEREFUELING

1. Boom in Trail
A. Extended 10 feet

*Ready for Contact

B. Fully extended

1. Tanker Manual Operation
without Tanker Disconnect
Capability
2. Acknowledge Receiver's
Manual Boom Latching Signal

C. Fully retracted

Offload Complete

2. Boom Stowed
A. Fully retracted

Tanker Air Refueling System
Inoperative

B. Extended 5 feet

System Malfunction, Tanker andReceiver Check Air Refueling
Systems

3. Flashing Receiver Director
Lights/Tanker Lower Rotat
ing Beacon ON

BREAKAWAY

4. **Receiver Director Lights
Going OUT During Contact

Tanker Request for Disconnect,
Receiver return to Precontact
Position

5. Receiver Closing and Opening Receptacle Door when in
Precontact Position

1. Manual Boom Latching

6. ***Steady Light from Receiv
er or rocking of wings

Emergency Fuel Shortage Exists

7. Flashing light from receiver
cockpit area

Initiate toboggan maneuver

2. Acknowledge Tanker's
Manual Operation without
Tanker Disconnect Capability
SignalFigure 83. (Sheet 1)

Change 187

T.O. GR1F16CJ1Visual Signals (Cont)
INDICATION
SIGNAL

8. (a). (DAY): Same receiver re
turns to precontact with re
ceptacle door open. Pilot signals closed fist, thumb tomouth, plus hand signaling
number****(NIGHT): Same receiver
returns to precontact with re
ceptacle door open, ready forcontact

BOOM AIR REFUELINGPROBE & DROGUEREFUELING

Additional fuel required
EMCON 24(b). (DAY): Same receiver re
turns to precontact, ready forcontact. Pilot signals closed
fist plus hand signaling number****(NIGHT): Same receiver
returns to precontact, readyfor contact

Additional fuel required
EMCON 24

*Receiver(s) in the observation position will move to the precontact position in their briefed sequence only
after insuring that the boom is in the ready for contact position and the preceding receiver has cleared thetanker. The receiver will stabilize in the precontact position, then move to the contact position. The boom
operator will not give the ready for contact signal until the preceding receiver has cleared the tanker.
**The receiver(s) will advise the tanker of any pilot director light malfunctions/deficiencies.
***If fuel shortage occurs at times other than scheduled air refueling, the receiver should be positioned so thesignal may be seen from the tanker cockpit.
****Additional fuel offloaded will be 5000 pounds for large receiver aircraft, or 2000 pounds for small receiveraircraft, on each subsequent contact.

Figure 83. (Sheet 2)88Change 1

T.O. GR1F16CJ1Exterior Lighting (KC-135)*A/R FLOOD LIGHTS*NAVIGATION/POSITIONSTROBE LIGHT* NACELLE ILLUMINATION LIGHT

LIGHT

BOOM MARKER
LIGHTS* BOOM NOZZLE LIGHT*NAVIGATION/POSITION
LIGHT*NAVIGATION/POSITION
LIGHT

LANDING LIGHT* UNDERWING ILLUMINATIONLIGHT

TAXI LIGHT* RECEIVER PILOT DIRECTOR
LIGHTSTAXI LIGHTSTROBE LIGHT*NAVIGATION/POSITION

LANDING LIGHT

LIGHT* UNDERBODY ILLUMINATIONLIGHTTERRAIN LIGHT(RETRACTABLE)* NAVIGATION/POSITION
LIGHT

VIEW LOOKING UP

NAVIGATION/POSITION
LIGHT*DESIGNATES ADJUSTABLE LIGHTING
1F-16X-1-8004X

Figure 84. (Sheet 1)

Change 189

T.O. GR1F16CJ1Exterior Lighting (KC-10/KDC-10)
FORMATION LIGHTSUPPER FUSELAGE
FLOODLIGHTSLOGOFORMATION LIGHT
ANTICOLLISION/RENDEZVOUSUARRSI FLOODNOZZLE

MARKERANTICOLLISION/RENDEZVOUS

HOSE FLOODLIGHT
FORMATION LIGHTSRECEIVER AIRPLANEFLOODLIGHTSWING TIP
ANTICOLLISIONFWD POSITION
WING TIP ANTICOLLISION

AFT POSITION

WING/ENGINE NACELLE SCANNING

HORIZONTALSTABILIZERUNDERSIDE LIGHTSAFT FUSELAGE

WING

DIRECTOR
LIGHTS

HOUSE SIGNAL
NOTE:
THE INTENSITY OF ALL
EXTERIOR LIGHTING IS
ADJUSTABLE.
LEFT SIDE SHOWNRIGHT SIDE SAMEFORWARDFUSELAGEUNDERBODY
LIGHTS
VIEW LOOKING UP

RECEIVERAIRPLANEFLOODLIGHTS

1F-16X-1-8003X

Figure 84. (Sheet 2)810Change 1

T.O. GR1F16CJ1
KC135 aircraft in cell formation will set their strobe
lights according to their position in the refueling cell.
Their lighting is as follows:
D Tanker Number 1

Upper and Lower Lights - RED

D Tanker Number 2Upper and Lower Lights - WHITED Tanker Number 3Upper Lights - RED
Lower Lights - WHITED Tanker Number 4Upper Lights - WHITE
Lower Lights - RED

Receiver Director Lights KC135
Refer to figure 85. The receiver director lights do notgive true vertical and horizontal information. The up
and down lights change because of angularmovement of the boom and the foreandaft lights
change because of inandout movements of the boom.
The axis of the director lights system is inclined at a
30 degree angle to the fuselage. This angle causes aninteraction in both lights when a true vertical or
horizontal movement is made by the receiver. For
example, flying straight forward while in contact willcause the boom to compress and also increase its
angle with the tanker fuselage. The lights will showthat the aircraft is flying forward and down. If a true
up movement is made, the boom will both compress
and lessen its angle with the tanker fuselage and the
director lights will indicate that an up and forward
movement has been made. Small fore and aftcorrections can be made with little or no power change
by moving vertically.
Receiver director lights are on the bottom of thefuselage directly aft of the nose landing gear. They
consist of two rows of lights: the left row for elevationand the right row for telescoping. The elevation lights
consist of five colored panels with a green stripe,
green and red colors, and two illuminated letters, D
and U, for down and up respectively. The coloredpanels are illuminated by lights that are controlled byboom elevation during contact. There is an
illuminated white panel between each panel to serve
as a reference. The letters A, for aft, and F, forforward, augment the colored panels on the telescopeside. The receiver pilot director lights will remain
illuminated and follow boom movements in both thecontact made and disconnect conditions. There are nolights for azimuth position. A fluorescent yellow strip
on the bottom center of the tanker fuselage may be

used as centerline reference by the pilot. The
triangularshaped panels are for elevation and therectangularshaped panels are for forward and backmovement.
Receiver Director Lights KC10/KDC10
Refer to figure 85. The receiver director lights consistof two rows of lights located forward of the wing root.
Relative elevation position is provided by the left row
and the right row provides telescoping position. The
elevation row contains one striped green, two amber
and two red triangular panels, and two white letters:
U at the forward end for UP, and D at the aft end for
DOWN. The colored panels and letters are dimly
illuminated by background lights. The telescoping
row contains one striped green, two amber, two red,and four white rectangular panels, and two whiteletters: A at the forward end for AFT, and F at the aftend for FORWARD. The colored panels are notbackground lighted; however, the letter at each end of
the row is dimly illuminated. Separation is provided
by the white panels. The pilot director lights areadjusted by the boom operator to the size air refuelingenvelope for each receiver and provide guidance
during contact.
To provide more response time, the appropriate panel
and letter are illuminated in anticipation of receiver
movement. The director lights provide commands
based on both receiver position and rate of movement.
Figure 85 shows the lights with no receiver motion.
With rapid motions of the receiver, the lights canshow a correction required even though the receiveris in the center of the envelope. The red panel andletter at the ends of each row can be illuminated by
the boom operator to aid the receiver in attaining thecontact position.
NAVIGATION AND POSITION REPORTING
When rendezvous is complete, tankers will be
responsible for all navigation, weather avoidance andposition reporting. The tanker pilot will, once eachhour, advise the receiver pilots of the cell geographic
position, heading, and distance and ETE to the next
checkpoint or destinations applicable. If the receiverpilot has not completed his onload upon reaching his
geographic air refueling abort point, the tanker pilot
will so advise.
NOTE
During refueling operations conductedwithin a preplanned orbit pattern, thetanker is not required to provide theabove information unless requested by
the receiver leader.
Change 1811

T.O. GR1F16CJ1
Post Air RefuelingUpon completion of air refueling, the tanker will
normally climb to the top and the receiver(s) will
descend to the bottom of the air refueling block. The
receivers should maneuver to the prescribed
formation position while awaiting post air refuelingreport and further clearance. The tanker will give
post air refueling information to the receiver as
required. The receiver will advise the tanker of any
pilot director light malfunctions/deficiencies, e.g.,lights intermittent, inoperative, dim, dirty, etc. Upon
termination of air refueling, all exterior lights will be
as directed.

Receiver will insure a safe clearance
from the tanker(s) as they proceed on
their assigned missions. Receiver(s)required to accelerate past the tanker(s)and climb on the refueling heading will
maneuver either left or right (aminimum of 1 NM) of track to preclude
climbing directly in front of thetanker(s) remaining receiver(s). Air
craft flying through departing receivers'
jet wash may experience damage to theaircraft and injury to personnel.

812Change 1

FLYING SAFETY
BOOM ENVELOPE LIMITS
Refer to figure 86. The refueling envelope is limitedby the refueling receptacle location. As long as thereceiver is positioned within these limits, contact canbe maintained despite rolling, yawing, or pitching.

FApproaching boom limits at a relatively

high velocity can cause structural
damage as a result of an inability to
disconnect due to binding of the boom
nozzle.

FDue to the restricted refueling envelope,boom limit switch protection is notprovided in up elevation.
NOTE
The boom operator will disconnect at 25
degrees elevation (upper envelope limit).
In this position, a green DOWN arrow
will be illuminated.

T.O. GR1F16CJ1

RECEIVER DIRECTOR LIGHTSILLUMINATION PROFILE (KC-135)

XX1F-16A-1-0162X

Figure 85. (Sheet 1)

Change 1813

T.O. GR1F16CJ1

RECEIVER DIRECTOR LIGHTS (KC-10/KDC-10)

XX1F-16A-1-0163X

Figure 85. (Sheet 2)814Change 1

T.O. GR1F16CJ1Boom Envelope Limits (KC-135)
RIGHT AZIMUTH LIMIT

1010LEFT AZIMUTH LIMIT

20

6 FT EXTENDED
INNER LIMITUP ELEVATION LIMITDOWN ELEVATION LIMIT
18 FT EXTENDED
OUTER LIMIT

40

1F-16X-1-8002X

Figure 86. (Sheet 1)

Change 1815

T.O. GR1F16CJ1Boom Envelope Limits (KC-10/KDC-10)
RIGHT AZIMUTH LIMIT

23

23

LEFT AZIMUTH LIMIT

25

UP ELEVATION LIMIT

6 FT EXTENDED
INNER LIMITDOWN ELEVATION LIMIT
21 FT EXTENDED
OUTER LIMIT

40

1F-16X-1-8001X

Figure 86. (Sheet 2)816Change 1

T.O. GR1F16CJ1

AIR REFUELING PROCEDURESNOTEDEPARTURE/ENROUTE PROCEDURES

On runways less than 300 feet wide, thereceiver will remain in the number 1
position until the tanker rolls.

Cell Leader Responsibility
The tanker cell leader or specified commander is
responsible for the command and control of theformation and the air refueling operation. The cell
leader or specified commander will coordinate withthe receiver force commander to insure successful
mission completion. Formation integrity and disci
pline begin with the formation briefing. The cell
leader must insure that all aspects of the mission areclarified and understood.
Normally, lead responsibilities pass from number one
to number two to number three; however, all pilots
must be prepared to assume full responsibility for theformation at any time. When it is necessary totransfer lead responsibilities, all airplanes in theaffected formation will be notified when the newleader assumes responsibility for the formation.
The tanker cell leader/specified commander musttake every feasible action to enhance the possibilityof completing air refueling.
Buddy Departure
A buddy departure is effected when the tanker(s) andreceiver(s) take off from the same base and visualcontact is maintained.

TAKEOFF

Wake turbulence generated by a preced
ing aircraft may create a hazard during
buddy takeoffs and joinups.
The tanker will roll first, followed in 45 seconds by the
first receiver in his element. Each element will be
individually cleared for takeoff by the tower after thelast aircraft in the preceding element has passed theend of the runway. Takeoff interval may be varied
when weather, terrain, airfield conditions, or otherlocal considerations dictate.
ABORTS DURING TAKEOFF
An aborting aircraft will make an abort call on theprebriefed common frequency as soon as possible.
Frequency changes will not be made by tanker/receiv
ers until all airplanes in the same element are
airborne.

TAXI
After engine start, receivers will check in with thetanker on the predetermined frequency. When readyto taxi, each tanker will call, (Tanker Call Sign),
Taxiing." A distance of 300 feet will be maintained
between tankers and receivers. Tankers not sched
uled to be used during the first air refueling will taxi
and takeoff first, followed by refueling elements
consisting of a tanker and his mated receivers for the
first refueling. Spare tankers will taxi and takeoff
last.
LINEUPOn runways at least 300 feet wide, the KC135 willline up on the downwind side of the runway. The
receivers will be positioned on the upwind side of the
runway, maintaining wing tip clearance.

ELEMENT JOINUP AND CLIMB
The climb speed schedule will be based on the firsttanker/receiver element climbing at 300 KCAS.
Individual tankers launching first will reduce climb
airspeed so that a 10 KCAS differential exists
between individual tankers and between the lastindividual tanker and the first refueling element.
Example:Tanker One . . . . . . . . . . . . . . . . . . . . . . .
Tanker Two . . . . . . . . . . . . . . . . . . . . . . .
Lead Element . . . . . . . . . . . . . . . . . . . . .
Second Element . . . . . . . . . . . . . . . . . . .
Third Element . . . . . . . . . . . . . . . . . . . .
Fourth Element . . . . . . . . . . . . . . . . . . .

280 KCAS
290 KCAS
300 KCAS
310 KCAS
320 KCAS
330 KCAS

Change 1817

T.O. GR1F16CJ1
NOTE
For two and three element cells, tankers
without receivers are considered elements. The lead element will climb at
300 KCAS and the following speed
schedule will be utilized to expedite
closure: In a two element cell, the second
element will climb at 330 KCAS. In athree element cell, the second and third
elements will climb at 320 KCAS and
330 KCAS, respectively.
If a ceiling is to be encountered prior to the completionof the joinup, the tanker will leveloff with aminimum of 500 feet clearance below the cloud layer
and maintain assigned climb airspeed to provide
visual flight conditions during the element joinup.
The air refueling element (tankers and receivers)joinup will be accomplished on the outside of the turnor on the tanker's left wing during a straight aheadjoinup. All turns by the tanker will normally be 20
degrees of bank (Exception: See Lost Wingman
Procedures). After the receiver flight has stabilized inits proper formation position, the element leader will
inform the tanker pilot.

altitude separation will be used for deconfliction. The
following procedures will be used to change positionsof flights in trail:1.

The tanker should be in straight and level
flight.

2.

The flight/element departing the tanker
reforms and moves back from the right wingwith offset and establishes contact (Radar,
Tacan, etc.) with the tanker. Maintain altitudedeconfliction from any joining receivers.

3.

The next flight/element may then begin to
move forward maintaining altitude.

4.

Once the flight/elements have horizontal
separation (established by Radar, Tacan, etc.),
the next receivers can climb and the departing
receivers can descend to briefed altitudes.

5.

If a flight/element is to rejoin from the trail
position and the preceding flight/element
elects to stay in formation with the tanker, theflight/element in trail will close on the tanker
until they are established in a visual position
and then switch positions with the precedingflight.

Force Extension Procedures
Element Cruise Formation
Receivers will fly route formation position on thetanker. (For extended VMC cell cruise formation,
receivers may fly a spread formation.) Route andspread formations will be as specified in theoperational procedures manuals. Spacing may be
reduced during IMC or night operations. When air
refueling is required, the other receiver(s) willassume the observation position.
At night or in IMC, a maximum of 12 aircraft will beassigned to one tanker - a maximum of three aircraft
will fly on each wing of the tanker, and additional
aircraft/elements will fly in a 13 NM trail position, aminimum of 1000 feet below their tanker's altitude.
During tanker cell operations, only the last tanker
should have receivers in trail. With nine or lessreceivers, only one element need be in trail (oneelement refueling while the other two elements are
switching positions). With 10 to 12 receivers, twoelements will be in trail (one element maintaining
trail, and the other two elements switchingpositions - and with four elements, one element
refueling). In this case, offset trail and additional

818Change 1

Force extension procedures are used when force
extensiontankersair
refueltankers(KC10/KDC10/KC135) escorting fighters during
deployment operations. Force Extension missions areoften complex and demanding to all aircrews,
especially in IMC. All facets of the mission, to include
the rendezvous, formation, air refueling, VMC/IMC
rejoin procedures, and cell breakup will be briefed
and clearly understood by all participants during
mission planning. Basic Point Parallel or Enroute
Rendezvous procedures apply, except as indicated.
Multiship tanker cells will fly 60 degree echelon, 2
NM spacing unless otherwise directed. Missions thatencounter IMC conditions during air refueling may
increase air refueling echelon formation spacing from
2 miles to 3 miles. If a midmission rendezvous isplanned, the escorting tanker will attempt to contact
the force extension tanker and pass the ETA to the
ARCP and updated weather information which may
effect air refueling. Inflight visibility will be thedetermining factor in utilizing VMC versus IMC
procedures to conduct air refueling. For the purposeof these procedures, VMC is defined as visibility equal
to or greater than 2 NM. IMC is defined as visibility
less than 2 NM.

T.O. GR1F16CJ1
VMC PROCEDURESFighters will join on the force extension tanker whencleared by the escorting tanker pilot. After all fighters
have joined on the force extension tanker, theescorting tanker will be cleared for refueling.
Fighters should fly a loose wing formation andremain with their force extension tankers in the eventa breakaway occurs.
IMC OR NIGHT PROCEDURES
Airtoair radar equipped fighters, when cleared by
the escorting tanker pilot, will fly 1 1/2 to 2 NM trail(6 o'clock) position, 2000 feet below their assigned
force extension tanker. Nonairtoair radar equippedfighters will comply with the VMC procedures.
POST REFUELING PROCEDURES
Once all air refuelings are complete, the escortingtanker(s) will descend 1000 feet, offset slightly to theright and then move to a position 1 NM in front of theforce extension tanker(s). Once the escortingtanker(s) are stabilized 1000 feet below and forwardof the force extension tanker, the escorting tanker willassume lead for the formation after a positive verbal
lead change. The escorting tanker will then clear thefighters forward to rejoin.
Nonairtoair radar equipped fighters will rejoin
visually with their respective escorting tanker.
Airtoair radar equipped fighters will rejoin using
radar guidance if required. If IMC prevails and poor
visibility precludes visual rejoins, force extension
tanker(s) may momentarily reduce separation to 1/2
NM and 500 feet vertical separation to facilitate therejoin.
Once all fighters have rejoined on their respective
escorting tanker, the force extension tanker(s) willdepart the stream from the rear of the formation.
If the formation reaches the End A/R point and visualrejoins are not possible, force extension tankers willcontinue along the receivers' route of flight until
visual rejoins are possible, fuel permitting. If theforce extension tanker(s) reach BINGO FUEL at, or
after, the End A/R point and the fighters have notrejoined with the escorting tanker(s), the entire
formation will abort to a suitable alternate airfield.
The tanker radar should be used for position
monitoring throughout the maneuver. It is the force
extension tanker leader's responsibility to inform theentire formation of current heading and airspeed
until relieved of that responsibility by a lead change.
The force extension tanker/cell will reform at the topof the air refueling block. Once the fighters haverejoined on their respective escorting tankers, theescorting tanker/cell will reform at the bottom of theblock. Cell separation will be accomplished by theforce extension or escorting tanker cell increasing/de
creasing airspeed as determined by mission require
ments and/or the premission brief. Force extension
or escorting tankers will not make any climbing ordescending turns to depart the stream until thetanker cells are identified visually or by radar, arewell clear, and verbal coordination is made between
tanker cell leaders. All aircrews must clear
aggressively and be cognizant of potential converging
headings or conflicts.
When simultaneous refueling of fighters andescorting tankers is required, 310 KCAS
(KC10/KDC10) or 295 KIAS (KC135) normally will
be used as refueling airspeed for the formation. The
lead force extension tanker will determine air
refueling airspeed based on aircraft type, altitude,
weight, weapons load, etc.
The above procedures are the standard for force
extension tankers refueling escort tankers. Any
deviations to these procedures are not authorized
unless coordinated between all tankers (escortingand force extension), receivers, and mission commander. If a predeparture briefing is not conducted
due to geographically separated departure locations,the escorting tanker will coordinate changes inflightto the force extension tanker upon initial contact,prior to the air refueling rendezvous.
Lost Wingman Procedures
In the event a receiver aircraft becomes lost during
refueling operations or during buddy cruise, thefollowing procedures will apply:1.

If a tanker or receiver detects weather (by radar,
reduced visibility, etc.) which may require a
receiver to go lost wingman, the tanker shouldmake any required navigational corrections,then fly straight and level. Any subsequent
turns required should be made with ten degrees
of bank maximum and called over the radio (i.e.,Tanker 21 is rolling into a ten degrees rightturn.").

2.

The receiver flying on the tanker's wing will
simultaneously inform the leader and turnaway using 15 degrees of bank for 15 seconds,
then resume heading.

3.

The receiver flying on the wing of the above
receiver (second in the echelon) will simultaneo
usly inform the leader and turn away using 30
degrees of bank for 30 seconds, then resume
heading.
Change 1819

T.O. GR1F16CJ1
4.

5.

The receiver flying on the wing of the above
receiver (third in the echelon) will simulta
neously inform the leader and turn away using
45 degrees of bank for 30 seconds, then resume
heading.

FFor receiver's directed rendezvous, theThe receiver who is flying in the precontact orcontact position will simultaneously reduce
airspeed 10 knots and descend 500 feet below
refueling altitude. Hold refueling heading andafter 30 seconds resume normal airspeed.

FGround radar assistance should be6.

If the above procedures are started while
tanker(s) are in a turn, request tanker(s) rollwingslevel and call rollout heading.

7.

Notify flight leader or tanker commander of thesituation.

8.

Attempt rejoin only after receiving clearance
from the tanker and when within radar or VMC
capability.

Abort Procedures
If a receiver aborts during an air refueling mission,
the receiver leader will determine the course of action
to be taken.
RENDEZVOUS PROCEDURES
NOTE
Initial visual contact between thereceiver and tanker may be enhanced,
inflight weather conditions permitting,if the tanker jettisons fuel to increase its
visual signature. This procedure may beinitiated/requested by the tanker, re
ceiver, or the ground agency controlling
the rendezvous. It should only be used if
a receiver low fuel state or other similar
circumstances require the rendezvous
be expedited.
The type of rendezvous will be dictated by mission
requirements, weather conditions, etc. If weather
conditions for joinup at cruise altitude are VMC, anenroute rendezvous may be used. When tanker andreceivers operate as separate flights, the Point
Parallel rendezvous will be primary.
After radio contact has been established, bothtankers and receivers should be tuned to the same
NAVAID, if possible, to improve rendezvous
capability.
820Change 1

NOTE
receiver's airborne radar equipment (ifavailable) will be the primary means toaccomplish a rendezvous. A/A Tacan
will be used as a backup.
utilized to the maximum for all
rendezvous. The tanker commander
will monitor ground control radio
frequency and confirm receiver range/
bearing information to insure positive
identification during the rendezvous.
Ground radar assistance will be termi
nated when the receiver reports visualcontact with the tanker.

Altimeter Settings
Unless otherwise directed, an altimeter setting of
29.92 inches Hg will be used for air refueling
operations at or above transition altitude or whenover water and operating in accordance with ICAO
procedures. For all other air refueling operations, thebriefed altimeter setting will be used.
Rendezvous Altitude Block
Four consecutive altitudes shall be requested by thetanker for rendezvous and refueling. When four
altitudes are available, the rendezvous will beeffected with the tanker at the second altitude andthe fighter(s) at the third. For example, when therefueling altitudes are FL290, 300, 310, and 320, thetanker will be at FL310 and the fighter(s) at FL300
for rendezvous, thus providing 1000 feet above thetanker and 1000 feet below the fighter(s). Whentankers are in cell, they will stack up from the second
altitude from the top of the block, FL310 in thisexample.
When only three altitudes are available, the tanker
shall be at the top altitude with the fighter(s) at the
middle altitude, providing 1000 feet below thefighter(s).
When only three altitudes are available and therefueling involves tanker cell formation, the highest
tanker within the cell should be at the top of the block.
To accomplish this, the tanker leader shall place
himself at an altitude that will permit the highest
tanker in his cell to be at the top of the block. For
example, when there are two tankers and theavailable block is FL270290, the tanker leader would
be at FL285 with number two tanker at FL290. The
fighters would rendezvous at FL275 (1000 feet below
the lowest tanker).

T.O. GR1F16CJ1
The above procedures do not apply when aircraft areoperating on an Altitude Reservation (ALTRV) or
when clearance has been granted for aircraft tooperate as an enroute cell. In these cases, the altitude
block will provide airspace necessary to accommodate
the type of formations being used (standard or
nonstandard), with at least 1000 feet between thehighest receiver and lowest tanker during rendez
vous, and at least 1000 feet below the air refuelingformation once the rendezvous is complete.
Track
Receivers shall pass over the ARIP, if applicable, andmake good the planned inbound track to the ARCP. If
a deviation is required because of weather, etc.,
receivers will not attempt rendezvous or proceed to
the ARCP until the deviation has been approved by
air traffic control and coordinated with the tanker. If
radio contact between the tankers and receivers is notestablished prior to the ARCT, the tankers will beover the ARCP at the ARCT.
Tanker Rendezvous Equipment
Tanker rendezvous equipment consists of thefollowing:
D KC135
S A/A TACAN  DME only
S Radarbeacon AN/APN69 (all aircraft) and
AN/APN134 (some aircraft)

S Automatic Direction Finder  AN/ARA25
D KC10/KDC10S A/A TACAN  Range and Bearing
S RadarWidth)

beacon APX78 (Two Pulse, Variable

S UHF/DF
S INS

NOTE
Radar beacon contact will be shortrange due to radar polarization incom
patibility.

Receiver Formation During Rendezvous
Formation procedures after leveloff or from the ARIP
until joinup with the tankers will be as follows:
NOTE
Formation lead changes and joinups
will normally be completed prior to
departure from the ARIP. Should suchmaneuvers be required subsequent to
departure and prior to joinup on thetanker(s), the rendezvous will not be
continued unless the flight leader ispositive of his position in relation to thetanker(s) and the published A/R track.
D Day VMC (Visibility five miles or greater). Flights

will be in trail, offset to the right of the precedingflight. When all aircraft are in visual contact withthe tanker(s), each aircraft/flight will join with his
respective tanker as briefed.

D IMC or Night. The receiver formation will be formed

into flight(s) of four aircraft in close or routeformation. Succeeding flights will be in a likeformation, positioned 1500 feet to the rear of the firstflight or 23 NM radar trail, depending on weather
conditions. When the tankers are established on theoncourse track, the receiver leader will position his
flight 3 NM to the rear of the last tanker. When alltankers in the cell are in positive radar contact,
receivers will climb to 1000 feet below base altitude.
If visual contact has been established, theflight/element leader will initiate joinup on the lasttanker. When the appropriate wingmen have visualcontact and are within 1 NM of their tanker, theleader will drop them off and proceed to the nexttanker. The receiver leader will continue as above
until all wingmen are on their appropriate tanker,
then join the lead tanker. If not in visual contact at
3 NM, the receiver flight/element leader will clear
the last receiver flight to join on the last tanker. The
last receiver flight will then turn to the right and joinon the last tanker while climbing to 500 feet below
base altitude, then maintain this altitude until
visual contact with the tanker is established. The
first receiver flight will join on the number 2 tanker,
maintaining 1000 feet below base altitude until
visual contact is established. Each receiver leader
will offset his tanker target 15 degrees to the left andclose at approximately 50 knots above the tanker
KCAS. Aircraft with operable airborne radar
equipment will close no closer than 1500 feet.
Receivers losing radar lock between 1 NM andminimum range will ensure 1000 feet altitude
separation is maintained and discontinue rendez
vous attempts until adequate range separation (1NM) is achieved or radar lockon is regained. Range
Change 1821

T.O. GR1F16CJ1
closure limitation for nonradar equipped receiver(s)or receiver(s) without radar lockon is 1 NM. Whenvisual contact with the tanker has been established,
the receiver element will form in the precontact
position, and the receiver flight leader will turn left,then right back to track heading and join on the leadtanker in the cell using the same procedures. The
flight leader's wingmen, after flight separation, will
echelon to the right. If visual contact with thetankers is not established, receivers will maintain 15
degrees offset, applicable altitude, and minimum
slant range until cell termination procedures areaccomplished.
Early Arrival
Once a joinup is initiated, and it is necessary for thejoining receivers to hold while waiting for a precedingflight to complete their operations, the joining flightwill join in a position 600800 feet out (laterally) fromthe receivers in the observation position if in VMC; or
fly offset laterally 13 NM in trail if in IMC, andmaintain 1000 feet below tanker base altitude. The
decision on which side to join will be based on the
direction of orbit of the tanker, departure intentionsof the receiver flight refueling, and the presence ofadditional holding receiver flights.

Joining flights of receivers should not
close astern of a tanker that is conducting
refueling operations with other receivers.
The wake turbulence generated by theseaircraft during departure/changing of
positions, if encountered, can result inloss of aircraft control.
In the event the receiver(s) arrives ahead of thetanker at the ARIP or ARCP point, the receiver will
orbit at an altitude that ensures at least 1000 feetseparation between tanker and receiver or any
elements of tanker and receiver cells. If receivers holdat the ARCP, they will normally enter a left handholding pattern using 2 minute legs at 1000 feetbelow air refueling altitude.
Tanker Identification
Tanker identification is critical in congested refuelingairspace. Available aids used in any combination
should be used to confirm tanker location/identifica
tion prior to and during the rendezvous. These aids
include ground radar, tanker/receiver radar, INS, A/A
TACAN, UHF/DF steers, common ground TACAN
stations, and radar beacons/IFF/SIF interrogation
systems for receivers so equipped.
822Change 1

NOTE
If the radar beacon is to be used topositively identify the tanker, thereceiver should attempt to identify thetanker rendezvous beacon as soon as
possible so that an alternate method of
identification may be used if unable toreceive the beacon. Only the tanker cell
leader or his designated alternate, if
applicable, will operate the rendezvous
beacon.
Point Parallel Rendezvous
The point parallel rendezvous is used to effect a rapidjoinup between the tanker and receiver withminimum receiver maneuvering. The tanker andreceiver approach on reciprocal headings offset, leftor right, a distance equal to the tanker turn diameter.
At a predetermined turn range, the tanker executes
a turn to the receiver heading to roll out
approximately 13 NM ahead of the receiver.
Normally, the tanker aircrew has responsibility forthe overall refueling operation and rendezvous andestablishes the offset and turn point. When tanker
systems are degraded, the situation dictates, or for
training, the Tactical Air Controller or receiver may
be responsible for the execution of the rendezvous.
Specific rendezvous responsibilities will be inaccordance with command guidance. Receivers will
monitor the air refueling frequency and attempt toestablish contact as soon as possible. If both tankers
and receivers are on a common radio frequency toobtain ground radar rendezvous assistance, thechange to air refueling frequency may be delayed
until positive radar/visual contact is established. If
under radar control, obtain bearing and distance tothe tanker prior to changing to air refuelingfrequency.
A successful point parallel rendezvous requires thetanker to maintain the proper offset and the receiverto fly the specified rendezvous track from the ARIP to
the ARCP. Emission Option 2 will be the normal
rendezvous and air refueling procedure. The receiverwill call 15 minutes prior to the ARCT and relay callsign, ETA (minutes early or late), and altitude. The
tanker will then confirm his call sign, air refueling
altitude and timing (minutes early or late) if it willaffect the rendezvous. If either the tanker or thereceiver is not on appropriate rendezvous altitude,
another radio call will be made when the proper
rendezvous altitudes are established. Receivers
departing the ARIP will maintain cruise speed at1000 feet below tanker base altitude until positive
radar or visual contact is made from in trail with theappropriate tanker(s). The receivers will proceed

T.O. GR1F16CJ1from the ARIP to the ARCP using all navigational
aids available to maintain the centerline. When it isdetermined the receiver is at or inside the ARIP, thetanker will turn to, or continue on, the reciprocal of
the receiver's inbound track and will establish andmaintain the proper offset until reaching the planned
turn range. However, receivers will not deviate fromthe ARIP/ARCP centerline unless directed to do so by
the tanker.
The tanker will adjust to appropriate air refuelingspeed when rolled out toward the ARCP. The tanker
INS/DNS will be the primary means of maintaining
the offset and the A/A TACAN will be primary forrange information. To provide A/A TACAN ranging,
the tanker and the receiver (1 airplane per cell) willtune the assigned A/A TACAN channels 15 minutes
prior to the ARCT. The receiver will set thenumerically lower A/A TACAN Ychannel and thetanker will set the higher Vchannel. A/A TACAN
should be left in A/A until the receiver reaches
precontact.
NOTE
Radio silence will be broken if thetanker or receiver determines thateither the tanker or receiver will exceed
ATC protected airspace while maneu
vering to attain the offset.
The tanker will note the receiver's distance when
halfway through the turn back to the ARCP. This isthe best time to determine if an overrun condition
exists and the best time for visual sighting. If an
overrun condition exists, appropriate action should
be taken. Tactical Air Controllers or receivers who
have established radar contact and positively
identified the tanker will offset the tanker in order toestablish a 26 degree relative bearing at 21 NM
range. The receiver/weapons controller will offset thetanker as follows:
APPROXIMATESLANT RANGE(NM)BEARING(DEGREES)

100
80
60
50
40
30
25
21
20
13

5
6
8
10
13
18
21
26
27
45

NOTE
FThe above slant range/bearing chart

applies only to the KC135 at approxi
mately 28,000 feet. It is based on 9.5
NM lateral offset with the tanker
using 30 degrees of bank during theturn to refueling heading. The
KC10/KDC10 normally uses 25 de
grees of bank or less during the turn torefueling heading and approximately
an 11.5 NM offset. Normally, the turnrange with the KC10/KDC10 will be
26 NM greater than the turn range
with the KC135.

FWhenmore than one tanker isinvolved, it may be necessary to add 1
NM to the turn range for each
additional tanker in the formation toensure receiver leader is 3 NM in trailof the appropriate tanker at rollout.

At the turn range, the tankers will turn to therefueling track and adjust to refueling formation. The
receiver will assume responsibility for closing on thetanker at the turn range. The last tanker in a cell willturn the radar beacon to operate, single code, onrollout to refueling heading at the completion of theturn. The receivers will normally be 3 NM in trail ofthe tanker(s). When rolled out toward the ARCP, thetanker will adjust to the appropriate air refuelingspeed.
Point Parallel Rendezvous With Tanker Escort
Refer to figure 87. The receivers will join on theescorting tanker in the briefed sequence. The orbitingand escorting tankers are responsible for effecting therendezvous.
An air refueling anchor is a lefthand racetrackpattern with legs separated by 20 NM and normallya minimum leg length of 50 NM. Tankers will adjustfrom enroute cell formation to the air refuelingformation of 20 degrees right echelon, 1 NM
nosetonose separation, stacked up at 500 footintervals during the final turn to the air refuelingtrack.
When the receivers are inbound, the rendezvous will
be directed by the Tactical Air Controller or thereceiver leader. The Tactical Air Controller or thereceiver leader (as appropriate) will determine thetype rendezvous to be made. The tanker will adjust to
Change 1823

T.O. GR1F16CJ1
refueling airspeed when directed by receiver leader.
Receivers will rendezvous 1000 feet below refuelingbase altitude until visual contact is established. In
the event Tactical Air Control radar is not available
to control anchor refueling operations, the following
alternate procedure will be used:

Receiver TurnOn Rendezvous

The tanker will establish a normal point parallel
rendezvous at the anchor point. Receiver flights willproceed to an anchor ARIP a minimum of 70 NM
upstream from the anchor point. Receivers willrendezvous 1000 feet below the refueling basealtitude until visual contact is established. Normal
point parallel rendezvous orbit (fighter) procedures
will be used for the rendezvous.

Rendezvous Overrun

NOTE
Unless directed otherwise by the tanker,
the receiver flight will accomplish a 360
degree left turn at the upstream end of
the anchor pattern to enable the tanker
to turn toward the receiver flight for
rendezvous.
After the receiver flight is joined up, the anchor
pattern will be used for refueling. If cleared by thetanker commander, subsequent receiver flights may
depart the ARIP prior to the previous receiver flights
departing the anchor area when the receivers havethe capability to assure safe aircraft separation andto join on the tanker using receiver turnonrendezvous procedures.
Ensure at least 1000 feet separation is provided
between each joining flight, and between the highest
flight and the lowest refueling element until visualcontact is established. Use of secondary frequency isrecommended. To preclude conflict with receivers
clearing the tanker or during a breakaway, ensure
fighters maintain adequate intrail spacing from therefueling formation.
NOTE
If subsequent flights are cleared todepart the ARIP and air refueling is
not complete prior to joinup, a fighter
turn on rendezvous should be made. If
this is not possible, these fightersshould hold at the ARIP until air
refueling is complete.
824Change 1

Receiver turnon rendezvous will be conducted inaccordance with the procedures established incommand guidance. Receivers will maintain required
vertical separation until visual contact is made withthe tanker(s).

In the event of an overrun by fighters, the receiver(s)will pass 1000 feet below the tanker to insure positive
vertical separation. The receiver will decelerate to
290 KCAS and maintain air refueling heading. The
tanker will accelerate to 355 KIAS (350 KCAS) or
Mach 0.90, whichever is lower, and maintain air
refueling heading. When the tanker is in positive
visual contact ahead of the receiver, the tanker will
decelerate to air refueling airspeed and normal
closure procedures will be employed to establish
contact.
Enroute Rendezvous
DEPARTURE AND CLIMB
The receiver departure time will be adjusted to place
him at altitude in trail of the tanker.
The tanker will level off, on course, at the
programmed cruise altitude and establish 260 KCAS
to permit receiver overtake. The receiver will level off,on course, 1000 feet below the tanker's base altitude
and establish a closing airspeed. All available
rendezvous aids will be used to effect tanker/receiverclosure until visual contact is made. Receivers willcall visual" when visual contact is established withthe tanker. Tankers will accelerate to enroute or
refueling airspeed at the direction of the receiver.
ENROUTE RENDEZVOUS PROCEDURES
An enroute rendezvous may be used when thetanker(s) and receiver(s) fly individual flight plans to
a common rendezvous point (RZ), where joinup isaccomplished, and continue enroute cell formation to
the ARCP.
These procedures may provide an orbit delay ortiming triangle enroute to the ARCP. It is not
appropriate to accomplish a point parallel rendez
vous at the RZ because the length of the orbit legs
cannot be extended. Tankers will depart the RZ tomake good the ARCT or the receiver's ETA to the
ARCP.
Either tanker(s) or receiver(s) may be scheduled toarrive at the RZ first, orbit if necessary, and thendepart at a preplanned time.

T.O. GR1F16CJ1Racetrack Patterns (Typical)

Anchor Pattern

INBOUND COURSEANCHOR POINT

Anchor Pattern (Alternate Procedure)

INBOUND COURSE
ARIPRECEIVERHOLDINGPOINT

TANKERORBITANCHOR POINT

Enroute Rendezvous (Typical)

TIMING TRIANGLES

1F-16X-1-8006X

Figure 87.

Change 1825

T.O. GR1F16CJ1
NOTE

When close interval stream operations
are being conducted, do not use orbit
delays to control timing.
The RZ will be located a minimum of 50 NM prior to
the ARIP/SD. Tracks from the ARIP/SD may beestablished from any direction and need notnecessarily be an extension of the air refueling track.
If orbit delays are required, they will be accomplishedby orbiting at the RZ point along an extension of the
track from the RZ to the ARIP/SD. Orbit in a
racetrack pattern using 30 degree banked turns anda maximum of 15 NM straight legs (unlessoperational directives specify longer straight legs).
Tanker(s) and receiver(s) will joinup at the RZ bycontrolling timing so they arrive at the RZ at the sametime. Timing to the RZ may be adjusted using
differential airspeeds, orbit delays or timing
triangles. If a planned orbit delay is used, receiver(s)and tanker(s) may accomplish joinup in the orbit.
Assigned altitudes at the RZ will provide at least 1000
feet separation between affected airplanes (highest
tanker and lowest receiver), with the receivers
normally at the highest altitude. If the receiver(s)planned level off altitude is within 30 minutes flying
time from the ARIP, the receiver(s) may level off
below the tanker and maintain an altitude which
provides a minimum of 1000 feet vertical separation
between the highest receiver(s) and the lowest
tanker(s).
Communications will be in accordance with specified
emission option. If radio contact between the airplanes
has not been established prior to the rendezvous
control time, or the adjusted rendezvous control time,
airplanes will maintain altitude and depart the RZ tocross the ARCP at the ARCT. Delays at the ARCP will
use normal orbit procedures unless otherwise directed.
If there is minimal separation between following
aircraft or cells using the same track, orbits at the
ARCP will require close coordination and a thoroughcrew briefing to ensure altitude separation.
When the aircraft or cells pass the ARIP/SD, thetanker(s) and receiver(s) will echelon and thereceiver(s) will begin descent to the base air refueling
altitude. Receiver(s) will descend to be at the basealtitude 80 NM prior to the ARCP. Tanker(s) will
maintain published buddy cruise KCAS and adjust to
air refueling airspeed crossing the ARCP.

826Change 1

For peacetime training missions, theARIP or the ARCP may be designatedas the RZ. In these cases, cells will
echelon and start descent at the baserefueling altitude as soon as practicalafter rendezvous completion.
If prebriefed, tanker(s) and receiver(s) may adjust to
air refueling airspeed and begin air refueling after
passing the RZ. Once departing the RZ/ARIP, thetanker(s) should fly centerline. The receiver is the
maneuvering aircraft. If the tanker is behind thereceiver, the tanker should accelerate and passslightly off the left wing of the receiver.
Alternate Rendezvous Procedures
Tanker and receiver crews must be prepared at all
times to accomplish the rendezvous using whatever
resources are available. When rendezvous equipment
is degraded, tankers and receivers will fly the sameprofiles as described in previous paragraphs. The
following are some suggested alternate rendezvous
procedures which should be used in any combination
to ensure a successful rendezvous:
NOTE
Initial visual contact between thereceiver and tanker may be enhanced,
inflight weather conditions permitting,if the tanker jettisons fuel to increase its
visual signature. This procedure may beinitiated/requested by the tanker, re
ceiver, or the ground agency controlling
the rendezvous. It should only be used if
a receiver low fuel state or other similar
circumstances require the rendezvous
be expedited. If required, the tanker will
dump fuel in 500  1000 pound incre
ments until positive visual contact canbe maintained.
D Radar/Rendezvous

Beacons. The receiver/tanker
beacons may be used for range and offset
information with suitably equipped airplanes.
Depending on equipment capability, one airplane
should maintain the planned outbound or inboundtrack while the other airplane maneuvers toestablish the planned offset. The tanker will clearly
establish which airplane will be maneuvering.

T.O. GR1F16CJ1D Common Ground Station. If A/A TACAN is notavailable, switching to a common ground TACAN/
VORTAC station for range information may benecessary. The final turn to refueling track is madewhen the DME difference equals proper turn range.

D UHF/DF. For DF steers, receivers will be requested

to use the MIC switch without talking. The receiverwill transmit on the air refueling frequency
approximately 10 seconds out of every 20 second
period, ending each transmission with the receiv
er's call sign. When the receiver position shows
proper turn range bearing (No Wind) from thetanker heading, the tanker will turn to the refuelingtrack. Notify the receiver when the turn is started.
At the receiver's request, the tanker will transmit
a homing signal.

D ETA. When adequate navigational check points areavailable, the tanker may adjust final orbit pattern
to arrive over the ARCP on the air refueling headingat the receiver(s) ETA to the ARCP.

D Ground Radar Assistance. Ground radar facilities

may be used for vector and separation advisories.
Ground radar assistance will be used to the
maximum when conducting rendezvous withsignificantly degraded equipment to ensure a
successful rendezvous.

Emission Option 3

S Inbound courses to the RZ/ARIP
S Radio silent termination time in the event of a

missed rendezvous

Missed Rendezvous Procedures
If contact is not established at the RZ/ARIP, thetanker will arrive at the ARCP at the ARCT. This
procedure begins when either aircraft arrives at the
ARCP and does not have visual contact with theother. In this case, a left hand orbit should be enteredand orbit controlled so as to be over the ARCP atintervals of every eight minutes (ARCT plus 8, plus
16, etc.). While in the orbit, every attempt should bemade to establish visual contact with the otheraircraft. The length of the delay and decision as tohow long to continue radio silence should bedetermined during mission planning/development
prior to flight.
SPECIAL AIR REFUELING PROCEDURES
The tanker boom is controlled by the boom operator
while the fuel transfer (pressure, flow, quantity, etc.)
is normally controlled by the tanker crew from thepilots' compartment. In IMC, when visibility is suchthat Lost Wingman Procedures may be necessary,receiver formations and the refueling sequence will
be structured so that no more than three aircraft areon each wing of the tanker.

The elimination of the 15 minute prior calls increase
the element of risk, and the following guidelines
should enhance safety considerations:Refueling Sequence

D Normally accomplish when clear of clouds.

Normally, the leader will proceed to the precontact
position. Number 2 will proceed to the lead element's
observation position. The second element will proceed
to an observation position on the tanker's opposite
wing. Each subsequent receiver will visually clear
and move from the observation position to theprecontact position. The refueling sequence will be
designated by the receiver leader. Each receiver, after
refueling is completed, will rejoin to an outside wing
position of his original element. When all receivers
have completed refueling, the receiver force will
rejoin to the left or right, as briefed, and slightly below
the tanker.

D If unable to remain clear of clouds, tanker(s) andreceiver(s) will immediately confirm altitudes.

D The receiver and tanker inbound courses to the

RZ/ARIP must be separated by a minimum of 30
degrees.

D The

receiver and tanker inbound legs to the
RZ/ARIP must be a minimum of 40 NM in length.

This type rendezvous should be an enrouterendezvous at the ARIP with both aircraft using thesame RZ time. The receiver should rendezvous 1000
feet below the tanker. An ETE from the ARIP to the
ARCP should be planned which permits an airspeed
which falls in the middle of the aircraft speed
performance envelope. It is essential that crews/planners coordinate certain items during mission
planning/development. Minimum items include:
S Rendezvous altitudes
S RZ time and ARCTFINGERTIP FORMATIONECHELON FORMATION (VMC ONLY)
Normally, the leader will proceed to the precontact
position. Number 2 will proceed to the observation
position with the remainder of the flight. Refuelingsequence will be as directed by the receiver leader.
Each receiver will visually clear and move from theobservation position to the precontact position. The
receivers, after refueling is completed, will rejoin inechelon formation on the tanker's opposite wing.
Change 1827

T.O. GR1F16CJ1
FUpwash and downwash effects may

Fuel Management
The ENG FEED switch should be in the NORM or
proper position, and the fuel distribution will bechecked within flight manual tolerances on the fuelquantity indicator prior to contact with the tanker.
The fuel system operation is automatic (fuel being
distributed to internal and external tanks simulta
neously).

NOTE

occur, drawing the aircraft together.
Low pressure areas created by anoverrunning receiver flying under thetanker will affect static ports, causing
possible erroneous airspeed and alti
tude indications to both aircraft. The
tanker autopilot altitude hold function
may sense the low pressure as aclimbing indication and initiate adescent into the lower aircraft.

Boom and Receptacle Procedures
NOTE

FIf a partial fuel load is onloaded, a fuelspread in excess of flight manual limits
should be anticipated.

F B Disconnect from the boom may occur

before all tanks are full if the externalfuel tank configuration consists of only
a centerline fuel tank. Such a disconnect
typically occurs when refueling with an
initial internal fuel load of 4000 pounds
or more and the centerline tank empty.
At disconnect, the aircraft total fuel may
be up to 1600 pounds less than full, withmany occurrences resulting in approxi
mately 1000 pounds less than full.

Precontact
All precontact air refueling checks will be completed in
the observation position or prior to reaching 1 NM in
trail, except for final exterior light adjustment. Afterthe receiver has stabilized in the precontact position,
the receiver will move to the contact position.

For night operations, prior to closing
for contact with the tanker, coordinate
with the boom operator on exterior
lighting to avoid impairing nightvision.
When cleared, move forward to the contact position
and the boom operator will make contact. The
receiver may request assistance from the boom
operator in obtaining and maintaining position.
From the precontact position, the receiver moves
slowly with a 23 knot closure until reaching thecontact position. When closing on the boom, constantcross reference between the boom and the tanker
fuselage will alleviate any tendency to chase"variations of boom trail position due to turbulence.
When stabilized in the contact position, maintain thisposition. The boom operator will then make thecontact.

FIf the receiver director lights fail to

illuminate when contact is established,
the receiver pilot will inform the boom
operator if he wishes to continue
refueling operations. If refueling iscontinued, verbal corrections from theboom operator may be requested.

FAttempts to affect a contact during loss
FThe

receiver will stabilize in theprecontact position and attain a zerorate of closure. If the receiver fails to
attain a stabilized position, or it
becomes apparent that a closure over
run will occur, a breakaway will be
initiated. Failure to initiate a breakaway under closure overrun conditions
can result in a midair collision.

828Change 1of any air refueling lighting that
results in less than desired illumina
tion will be at the discretion of theboom operator.

To maintain proper contact elevation and boom
extension, refer to the director lights located on thebottom of the fuselage of the tanker (See figure 85).
While in contact position, there is freedom in all threeaxes as depicted in figure 86.

T.O. GR1F16CJ1
If, for any reason, fuel is not transferring or istransferring at less than normal rate, the receiver pilot
will disconnect and monitor the aerial refueling statusindicator. The bottom lamp (DISC) lights amber whena disconnect has been accomplished. The system willautomatically reset to ready and the top lamp (RDY)
relights blue after a 3second delay. A second contact
may then be accomplished. If this does not resolve theproblem, the pilot may then disconnect, confirm
disconnect with the boom operator, and recycle thesystem by closing and opening the slipway door using
the AIR REFUEL switch.

FFollow the boom operator's instruction

explicitly. To reduce the probability ofboom strike after disconnect, it may benecessary to remain in a stabilized
position to allow for aerodynamic
fairing of the boom control surfaces.

Another feature of the KC10/KDC10 is the
Independent Disconnect System. This system allows
the boom operator to obtain a disconnect even when
the receiver's toggles remain in the latched position.
This system should be used in lieu of a Brute Forcedisconnect.

DISCONNECT KC135

Quick Flow Air Refueling Procedures

In the event of failure to obtain a contact and after
each disconnect, the receiver will move aft andstabilize in a position in trail of the boom or inprecontact position and await clearance from theboom operator to return to the contact position.

Refer to figure 88. Fighter type receivers may use
Quick Flow procedures to expedite air refueling
operations. Quick Flow allows receivers to minimize
refueling time with maximum fuel transfer. Quick
Flow may be used during day or night operations, in
VMC conditions only. If it appears that the flight may
encounter adverse weather conditions, standard IMC
procedures will be used. Coordination between
tanker(s) and receivers prior to initiation of Quick
Flow procedures is required. Air tasking guidance,
direct communication with the tanker unit, or adding
the term Quick Flow" to the initial radio call will
satisfy coordination requirements. Tanker lead is the
final authority for Quick Flow operations. Rightechelon formation is normally used for Quick Flow;however, variations are authorized with flight leadcoordination and tanker lead approval.

FRemainstabilized in the contact
position until visually confirming a
disconnect has been made. This will
prevent damage to the boom and/or
receptacle through a brute force
disconnect.

FBrute force disconnects can occur

unintentionally as the result of rapidly
exceeding boom limits or failure of thereceptacle toggles to release when adisconnect is initiated.

DISCONNECT KC10/KDC10
The KC10/KDC10 aerial refueling boom is con
trolled by a digital flybywire system. Certain failure
conditions of this system may cause one or more axesof the boom control system to become inoperative.
Should this occur, the boom operator may not be able
to maneuver the boom to avoid striking the receiver
airplane. In this situation, the boom operator will
issue instruction to direct the receiver to a position
where a safe disconnect can be effected.

FWhen notified that a KC10/KDC10boom flight control system failure has
occurred, do not initiate a disconnect
unless directed by the boom operator.

Normally, the receiver flight will join on the tanker
with the flight lead moving to the precontact position.
Remaining aircraft will proceed to the rightobservation position. Once the flight lead commences
refueling, the second aircraft in the air refuelingsequence will move to the OnDeck position. The
OnDeck position is normally flown as a routeformation with approximately 10' spacing. When theflight lead completes refueling, that aircraft moves toan observation position on the tanker's left wing. The
second receiver moves from the OnDeck position to
the precontact and contact position. With three ormore receivers, the third receiver moves to theOnDeck position. The right to left flow continues
until all fighters have refueled. When the air
refueling operation is complete, the flight may departthe tanker or, if additional refueling is required,remain in echelon formation on the tanker's left wing
and reverse the Quick Flow procedures, with a left toright flow. The second receiver will assume a left
OnDeck position and Quick Flow will continue inorder. Additional receivers arriving prior to the firstflight completing refueling operations will remain intrail position until they are cleared by the tanker tothe observation and/or precontact position.
Change 1829

T.O. GR1F16CJ1

Quick Flow Air Refueling

1F-16X-1-8007X

Figure 88.

830Change 1

T.O. GR1F16CJ1
In the event of a breakaway, the OnDeck receiver
follows the receiver that was on the boom. Any
receivers on the wing will remain with the tanker. In
the event a breakaway is initiated while a receiver istransitioning from the observation position to theOnDeck position, that receiver will follow thereceiver that was on the boom.
Toboggan
When altitude and atmospheric conditions result inthrust requirements that exceed the receiver's
available thrust, a toboggan will be necessary. The
toboggan technique is a coordinated effort between
the tanker pilot and the receiver pilot in which
refueling is accomplished in a slight descent, allowing
the receiver to perform the refueling with available
thrust.
The receiver pilot must signal or call on boom
interphone that a toboggan maneuver will be
required before reaching full military power.
The tanker pilot will very gently reduce power andinitiate a rate of descent of approximately 300 FPM
while maintaining the air refueling airspeed
throughout the toboggan maneuver.

SEPARATION FROM A SINGLE TANKER
The tanker and receiver leader will coordinate on themethod of separation. Normally, after the receiver
flight has reformed, they will clear the tanker bydescending or as directed by the controlling agency.
The tanker will advise and receive clearance from thereceiver leader before changing altitude or heading.
Receivers will maintain a safe clearance from thetanker as they proceed on their assigned mission.
CRUISE CELL TERMINATION (VMC)
When cleared by the receiver leader, elements will
join on their respective flight leaders. The receiverforce will then reform to the left and slightly below
the lead tanker. After receiving clearance from thetanker leader and the appropriate controlling
agency, the receivers will proceed on their assigned
missions, maintaining safe clearance from thetanker formation.
CRUISE CELL TERMINATION (IMC)

Afterburner Air Refueling is not recommended.

Refer to figure 89. Ten minutes prior to reaching thecell termination point (if the point is other thandestination approach fix), the receivers will reform inleft echelon on the left wing of the tanker(s). Uponreaching the cell termination point, the tanker(s) will
climb straight ahead 3000 feet and then turn to thedesired track, maintaining cell formation. Receivers
will maintain heading, altitude, and airspeed for 3
minutes. At this time, if flight formation rejoin is
impractical, number 1 receiver element will descend
1000 feet below base altitude, number 2 receiver
element will descend 500 feet to the base altitude,number 3 receiver element will maintain altitude(which will be 1000 feet above the base altitude), andnumber 4 receiver element will climb 500 feet to an
altitude which is 2000 feet above base altitude.
Receiver elements will then proceed with their
mission independently.

Separation/Termination Procedures

CELL TERMINATION AT TERMINAL APPROACH FIX

Weather Abort Procedures
Receivers must take every feasible action to enhance
the possibility of completing air refueling. Suchactions include altitude and course deviations
necessary to avoid severe weather. Deviations, whenrequired, must be made judiciously. When thereceiver leader determines that weather conditions
are such as to make formation refueling hazardous,
he may abort the cell. When the cell is to be aborted,
the receiver leader will instruct the tanker leader toclear refueling track. Normal end refueling procedures will apply.
Afterburner Air Refueling

Following completion of air refueling, the receiver(s)will maneuver to the prescribed formation position,
obtain tanker post air refueling report, and return tothe primary refueling frequency (if applicable). Afterthe receivers have reformed, the tanker leader will
provide the receiver leader with present position in
relation to the planned completion point. Additional
information will be provided if requested; i.e.,weather information, nearest abort bases, etc. The
receiver leader will request the no wind heading anddistance to the next checkpoint unless he has apositive fix from which to navigate.

Due to the many possible combinations of tanker/re
ceiver formations, terminal destination weather, andterminal airfield penetration facilities, it is impracti
cal to designate one optimum method for penetration
at the destination. The following methods should beapplied as applicable:
1.

From the final air refueling point, tankers andreceivers can be scheduled at their individual
optimum airspeeds to provide spacing for the
penetration.
Change 1831

T.O. GR1F16CJ1
2.

After the receivers have a positive TACAN
lockon, they will normally depart the tankers
and proceed to destination as directed by theappropriate controlling agency.

3.

When available, approach control should be
used with enroute descents to obtain aircraft
separation.

4.

The element (one tanker/two receivers) may
penetrate as a unit. Weather minimums forthis type approach are 2500 feet and 3 NM.

ELEMENT PENETRATION
If conditions exist which necessitate a more
expeditious recovery (fuel shortage, emergency, etc.),
a basic cell penetration may be made. Penetration
airspeed and descent rate will be coordinatedbetween the tanker and receiver leader. When VFR,
the receivers will break off and enter initial for a VFR
landing.
RECEIVER RADAR REJOIN PROCEDURES (IMC)
If receiver radar rejoin is desired at the completion of
the cell termination, the following procedure will be
initiated:

NOTE

1.

Each receiver will maintain his respective
altitude.

A low fuel altitude will be designated2000 feet below base altitude for
immediate descent of receivers withlow fuel or an emergency condition.
Receiver altitude changes will becoordinated by the receiver cell leader
with the ATC agency. All aircraft of thecell will note individual altimeter
errors at the cruise altitude with 29.92
inches Hg set on the altimeter and fly
their assigned altitudes after cell
separation with these errors applied.

2.

The receiver formation leader will maintainheading and each of the following receiver
elements will simultaneously turn left 15
degrees on the formation leader's command.

3.

Numbers 2, 3, and 4 elements will maintainthis heading for 1, 2, and 3 minutes
respectively, and then resume the original
heading.

4.

The flight will then rejoin on radar using theprocedures for radar joinup with tankers.

832Change 1

T.O. GR1F16CJ1Cell Termination Procedure

30,000NO. 3

28,000NO. 2

BASE ALTITUDE

500 FTNO. 1

27,000

ALTITUDE - FEET

29,0001000 FT
26,000LOW FUELALTITUDE

25,000

ALTITUDES LISTED ARE FOR SAMPLE ONLY USING 27,000 FT AS A BASE ALTITUDE.

1F-16X-1-8005X

Figure 89.

Change 1833

T.O. GR1F16CJ1NORMAL AIR REFUELING PROCEDURES
Armament Safety Check
Prior to closing within lethal range of the tankers,
complete the following checks:
1. MASTER ARM switch-OFF or SIMULATE.
2. LASER ARM switch - OFF.

If making an outer limit disconnect,high separation rates should be
avoided to prevent damage to the boomor receptacle.
2. AR status indicator light - DISC.

3. SMS - Confirm ordnance safe.
4. CMDS switches (9) - As required.
Precontact
Prior to air refueling, the following checks will becompleted:
1. TACAN - As required.
2. Emitters (ECM/FCR/RDR ALT) - As required(Quiet/Silent/STBY/OFF).

Remain stabilized in the contact
position until positive visual confirma
tion of boom separation is confirmed by
the boom operator.
Post Air Refueling
1. AIR REFUEL switch - CLOSE.

3. HOT MIC CIPHER switch - HOT MIC.
4. Exterior lights (Night) - DIM, STEADY.
5. ANTI COLLISION light switch (Night) - OFF.
6. AIR REFUEL switch - OPEN.
7. AR status indicator light - RDY.
Contact
1. AR status indicator light - AR/NWS.
NOTE
Once contact is made, boom inter
phone communications can be estab
lished with the boom operator if HOT
MIC is selected. Volume is controlled
by the intercom volume control. Theboom interphone capability is provided on all KC10/KDC10 andKC135 tankers.
2. Fuel transfer - Monitor.
Disconnect
1. A/R DISC button - Depress momentarily,then release.

834Change 1

Failure to close the air refueling switch
will result in the FLCS remaining intakeoff and landing gains, the roll rate
restricted to a fixed value, and thefailure of external fuel to transfer.
2. AR status indicator lights (3) - Off.
3. Fuel Quantity - Check.
4. MASTER ARM switch - As required.
5. SMS - As required.
6. CMDS switches (9) - As required.
7. TACAN - As required.
8. FCR - As required.
9. RDR ALT - As required.
10. LASER ARM switch - As required.
11. Exterior lights - As required.

T.O. GR1F16CJ1EMERGENCY AIR REFUELING PROCEDURES
Breakaway Procedures
Relative position of both airplanes must be closely
monitored by all crew members during all phases of
air refueling. When either a tanker or receivercrewmember determines that an abnormal condition
exists which requires an immediate separation of the
airplanes, that crewmember will transmit the
breakaway call on air refueling frequency. Abnormal
conditions include excessive rate of closure, closure
overrun, and engine failure. The receiver does nothave to be in the contact position to call a breakaway.
For all breakaways, transmit the tanker's call signand the word breakaway" three times (Example:
Chevy 2, breakaway, breakaway, breakaway") andsimultaneously take the following actions:
D Actuate disconnect switches as applicable.
D Retard throttle and establish a definite rate ofdescent, using speed break if necessary.

D If possible, drop aft of tanker until entire tanker isin sight and monitor flight instruments.

The tanker pilot will increase power to obtain forward
separation. Unless lateral separation cannot be
assured, the tanker will accelerate in level flight andwill not climb. The lower rotating beacon will beturned on, the pilot director lights will be flashed, andthe Radar/Rendezvous Beacon will be turned tooperate, if appropriate. When the receiver is wellclear, the breakaway may be terminated. The receiverpilot will be notified of and will acknowledge any
reduction in power by the tanker to resume air
refueling speed. If a climb is required, the tanker pilot
will disengage the autopilot and climb straightahead. If in a turn, the tanker will maintain theestablished bank angle until the receiver is well clear.
NOTE
FIf a breakaway is called prior to any

receiver reaching the observation
position, the entire receiver flight will
execute the breakaway procedure. If a
breakaway is called after receiver(s)have reached the observation position,
only the receiver in the contact orprecontact position will execute thebreakaway procedure. The receiver(s)in the observation position will maintain formation on the tanker.

FWith certain gross weights and aircraft configurations, the tanker rate ofacceleration on a breakaway may
exceed the rate of acceleration for thereceiver aircraft in the observation
position.

System Malfunctions
When any system malfunction or condition exists
which could jeopardize safety, air refueling will not be
accomplished except during fuel emergencies or whencontinuance of fueling is dictated by operational
necessity. At any time fuel siphoning is noticed, fueltransfer will be stopped and the receiver notified. The
requirement to continue fuel transfer will be at thediscretion of the receiver pilot.
NOTE
A small amount of fuel spray from thenozzle and receptacle during fueltransfer does not require fuel transfer tobe terminated. The receiver pilot should
be notified if this condition exists andthe air refueling operations will be
continued or discontinued at his discre
tion.
SLIPWAY DOOR WILL NOT OPEN
No backup system is provided to open or close theslipway door if hydraulic system B fails.
SLIPWAY DOOR WILL NOT CLOSE
If the slipway door will not close, perform thefollowing:
1. AR switch-CLOSE.
Normal FLCS gains and tank pressures will
be regained.
NOTE
The RDY, AR/NWS, and DISC lights
will not indicate normally. The NWS
light will not illuminate when nose
wheel steering is engaged.

Change 1835

T.O. GR1F16CJ1INOPERATIVE BOOM/RECEPTACLE LATCHING

Brute Force Disconnect

When all other recognized means of fuel transfer havefailed, and an actual fuel shortage emergency aboard
the receiver airplane exists, fuel can be transferred bymaintaining boom/receptacle contact using a slight
extend pressure on the boom telescope lever. Unusualand varying trim changes may be required of bothtanker and receiver airplanes.

There are two types of brute force disconnects:
Inadvertent and Controlled Tension.

If a fuel shortage emergency requires:
1. Boom operator-Inform of the need toaccomplish manual boom/receptacle pres
sure refueling.

NOTE
Enter any brute force disconnect as adiscrepancy in the AFTO Form 781.
The entry will specify which type ofbrute force disconnect occurred.
INADVERTENT DISCONNECT
An inadvertent brute force disconnect is defined as
any unplanned disconnect which is the result of oneof the following:
D The receiver aircraft moving rapidly to the aft limit,

causing mechanical tanker/receiver separation.

The receiver pilot must inform thetanker he is ready to receive fuel andcoordinate the disconnect cycle for theconclusion of refueling.

Prior to attempting this method oftransferring fuel, the boom operator
will brief the receiver pilot andthoroughly coordinate the procedures
to be used. Both tanker and receivercrews will monitor the refueling withextreme caution.
KC10/KDC10 Boom FLCS Failure
Do not disconnect until cleared by boom operator.

D Boom pullout occurs at 38 degrees elevation or

below.

Following an inadvertent brute force
disconnect, air refueling will be termi
nated except during fuel emergencies
or when continuation of air refueling is
dictated by operational necessity.
CONTROLLED TENSION DISCONNECT
A controlled tension brute force disconnect is defined
as an intentional, coordinated disconnect occurring
above 38 degrees elevation, accomplished by gradualmovement of the receiver aircraft to the aft limit, andending with a smooth tension boom pullout.
Coordination between the receiver pilot and the boom
operator is required to ensure as smooth a disconnect
as possible.
1. Slide out boom with gradual power reduction.

FWhen notified that a KC10/KDC10boom flight control system failure has
occurred, do not initiate a disconnect
unless directed by the boom operator.

FFollow the boom operator's instruction

explicitly. To reduce the probability ofboom strike after disconnect, it may benecessary to remain in a stabilized
position to allow for aerodynamic
fairing of the boom control surfaces.

836Change 1

2. When at full boom extension, tension
disconnect will occur with slight power
reduction.

A controlled tension brute force discon
nect will be accomplished only as a lastresort, after all other normal andemergency methods of disconnect havefailed.

T.O. GR1F16CJ1FFly stabilized at contact altitude until

certain the nozzle is clear of thereceptacle and slipway.

FThe receiver pilot must not jerk theboom out with rapid thrust change
toward idle or by using speed brakes; todo so may cause serious structural
damage. Gradual power reduction will
suffice to effect a disconnect.

FAir refueling for the receiver which

required controlled tension disconnect
will be terminated except during fuelemergencies or when continuation of
air refueling is dictated by operational
necessity.

Change 1837/(838 blank)

T.O. GR1F16CJ1GLOSSARY
STANDARD AND NONSTANDARD ABBREVIATIONSAAA
A/A, AA
AAM
AB
ac, ACA/C GW
ACM
ACMI
ADG
ADI
AFC
AGL
AGM
AIFF
AIM
AIS
AJ
AL
ALOW
ALT
AM
AMMO
AMRAAM
ANT
AOA
AOS
AP
AR, A/R
ARCP
ARCT
ARI
ARIP
ARMT
ASE
ASHM
ASIU
ASPIS
ATT, ATTD
AUX
AVTR

Advance Attack
Air to Air
AirtoAir Missile
Afterburner
Alternating Current
Aircraft Gross Weight
Air Combat Maneuvering
Air Combat Maneuvering
Instrumentation
Accessory Drive Gearbox
Attitude Director Indicator
Afterburner Fuel Control
Above Ground Level
AirtoGround Missile
Advanced Identification Friend or
Foe
Air Intercept Missile
Aircraft Instrumentation System
Antijamming
Aft/Left
Automatic Low Altitude Warning
Altitude or Altimeter or Alternate
Amplitude Modulation
Ammunition
Advanced Medium Range Airto
Air Missile
AntennaAngle of Attack
Angle of Sideslip
Autopilot
Air Refueling
Air Refueling Control Point
Air Refueling Control Time
AileronRudder Interconnect
Air Refueling Initial Point
Armament
Aeroservoelastic
Aft Seat HUD Monitor
Aft Station Interface Unit
Airborne Self Protection Integrated
SystemAttitude
Auxiliary
Airborne Videotape Recorder
B

BAK
BARO

Arresting Cable Prefix (e.g.,
BAK9)
Barometric

BATT
BDU
BIT
BL
BLU
BOS
BSU

Battery
Bomb Dummy Unit
BuiltIn Test or Binary Digit
Buttock Line
Bomb Live Unit
Backup Oxygen Supply
Bomb Stabilizing Unit
C

CADC
CARA
CBU
CCIP
CCM
CCRP
CCW
CDI
CENC
CFT
CG
CHAN
CIU
CIVV'sCNI
C&I
CONFIG
CONT
CO2
CRS
CRV
CSD
CSFDR
CTVS
CW

Central Air Data Computer
Combined Altitude Radar
Altimeter
Cluster Bomb unit
Continuously Computed Impact
Point
Coil Current Monitor
Continuously Computed Release
Point
Counterclockwise
Course Deviation Indicator
Convergent Exhaust Nozzle
Control
Conformal Fuel Tank
Center of Gravity
ChannelCentral Interface Unit
Compressor Inlet Variable Vanes
Communications, Navigation, andIFF
Communications and IFF
Configuration
Control
Carbon Dioxide
CourseCanadian Rocket Vehicle
ConstantSpeed Drive
Crash Survivable Flight Data
RecorderCockpit Television Sensor
Clockwise
D

dBA
DBU
dc, DC
DEEC
DED
DEEU
DEGR
DI

Adjusted (Human Ear Response)
Decibels
Digital Backup
Direct Current
Digital Electronic Engine Control
Data Entry Display
Data Entry Electronic Unit
DegradedDrag Index
Change 1Glossary 1

T.O. GR1F16CJ1
DIA
DIFF
DIS
DISC
DME
DN
DOI
DSG RS
DTC
DTOS
DTU
DTS
DVAL
DWAT

Diameter
Differential
DisableDisconnect
Distance Measuring Equipment
Down
Display of Interest
Designate/Return to Search
Data Transfer Cartridge
Dive Toss
Data Transfer Unit
Digital Terrain System
DValue
Descent Warning After Takeoff
E

EAS
ECM
ECP
ECS
EDU
EED
EGI
EGT
ELECT
ELEV
EMCON
EMER
EMS
EMSC
ENDUR
ENG
EOS
EPU
EQUIP
EST
ETA
ETE
EXT

Equivalent Airspeed
Electronic Countermeasures
Engineering Change Proposal
Environmental Control System
Engine Diagnostic Unit
Electroexplosive Device
Embedded GPS/INS
Exhaust Gas Temperature
Electronic (primary altimeter
operating mode)
Elevation
Emission Control Option
Emergency
Engine Monitoring System
Engine Monitoring System
ComputerEndurance
EngineEmergency Oxygen Supply
Emergency Power Unit
Equipment
Estimate
Estimated Time of Arrival
Estimated Time Enroute
ExternalF

FACK
FAULT ACK
FC
FCC
FCR
FFAR
FFP
FLCC
FLCP
FLCS
FLIR
FM
FO
FOD

Fault Acknowledge
Fault Acknowledge
Flight Control
Fire Control ComputerFire Control Radar
Folding Fin Aircraft Rocket
Fuel Flow Proportioner
Flight Control Computer
Flight Control Panel
Flight Control System
Forward Looking Infrared Radar
Frequency Modulation
Foldout
Foreign Object Damage

Glossary 2Change 1

FORM
FOV
fpm, FPM
FR
FT, ftFTIT
FWD

Formation
Field of View
Feet per Minute/Flightpath
MarkerForward/Right
Feet
Fan Turbine Inlet Temperature
Forward
G

g, G
GAAF
gal, GAL
GBU
GCA
GCAS
GCU
GD
GEN
GM
GND
GP
GPS
GRDCUSGS
GW

Force of Gravity
Ground Avoidance Advisory Func
tion
GallonGuided Bomb Unit
Ground Controlled Approach
Ground Collision Avoidance System
Generator Control Unit
Guard
Generator
Ground Map
GroundGroup
Global Positioning System
Gulf Range Drone Control Upgrade
SystemGlide Slope
Gross WeightH

HDG
HDG SEL
HF
HMCS
HQ
HSI
HUD
HYB VSVHYB
HYDRAZN
Hz

HeadingHeading SelectHigh Frequency
Helmet Mounted Cueing System
HAVE QUICK
Horizontal Situation Indicator
HeadUp Display
Hybrid Variable Stator Vanes
HybridHydrazine
Hertz
I

IAS
IAW
ICP
ID
IFF
IGV
IKP
ILS
IMC
IMSP
In., IN.
INC
IND
INOP

Indicated Airspeed
In Accordance With
Integrated Control Panel
Identification
Identification, Friend or Foe
Inlet Guide Vanes
Integrated Keyboard Panel
Instrument Landing System
Instrument Meteorological
Conditions
Instrument Mode Select Panel
InchesIncrease
Indicator
Inoperative

T.O. GR1F16CJ1INS
INT
INST, INSTR
INU
I/P
ISA

Inertial Navigation Set (or System)
Intensity or Internal or Interval
Instrument
Inertial Navigation Unit
Identification of Position
Integrated ServoactuatorJ

JETT, JTSN
JFS
JHMCS
JOAP

JettisonJet Fuel Starter
Joint Helmet Mounted Cueing
SystemJoint Oil Analysis Program
K

K
KCAS
KEAS
KIAS
KT(S)
KTAS
KVA

Thousand (e.g., 40K = 40,000)
Knots Calibrated Airspeed
Knots Equivalent Airspeed
Knots Indicated Airspeed
Knot(s)
Knots True Airspeed
Kilovolt Ampere
L

L
LADD
LAU
lb, LB
LB/HR
LB/MIN
LCO
LCOS
LD
LE
LEF's
LG
LMLG
LOD
LOX
LPU
LRU
LTS
LWD

Left
Low Angle Drogue Delivery
Launcher Armament Unit
Pound(s)Pounds per Hour
Pounds per Minute
Limit Cycle Oscillation
Lead Computing Optical Sight
Load or Low Drag
Leading Edge
Leading Edge Flaps
Landing Gear
Left Main Landing Gear
LightOff Detector
Liquid Oxygen
Life Preserver Unit
Line Replaceable Unit
LightsLeft Wing Down
M

M
MAAS
MAC
MAL
MAL & IND
MAN.
MAU
MAX
MAX AB
mb
MDTC

Mach
Mobile Aircraft Arrestment System
Mean Aerodynamic Chord
Malfunction
Malfunction and Indicator
ManualMiscellaneous Armament Unit
Maximum
Maximum AfterburnerMillibar
Mega Data Transfer Cartridge

MEC
MECH
MEM
MFC
MFD
MFDS
MFL
MFP
MHz
MIC
MIL
MIN
MK
MLG
mm
MMC
MPO
MRK BCNms
MSL
MUX BUS

Main Engine Control
Mechanical
MemoryMain Fuel Control
Multifunction Display
Multifunction Display Set
Maintenance Fault List
Main Fuel Pump
Megahertz
Microphone
Military
Minute or MinimumMark (equivalent of model)
Main Landing Gear
Millimeter
Modular Mission Computer
Manual Pitch OverrideMarker Beacon
Milliseconds
Missile or Mean Sea Level
Multiplex Bus
N

NA
NAM
NFOV
NLG
nm, NM
No., NO.
NORM
NOZ POS
NVG
NVIS
NWS

Not Applicable
Nautical Air Miles
Narrow Field of View
Nose Landing Gear
Nautical Miles
NumberNormalNozzle Position
Night Vision Goggles
Night Vision Imaging System
Nosewheel Steering
O

OAT
OBOGS
OCSC
OHEAT
OP
OPT
OSB
OVRD
OXY
O2

Outside Air Temperature
Onboard Oxygen Generating
SystemOvercurrent Sensing Contactor
Overheat
Operational or Optimum
Optional
Option Select Button
Override
OxygenOxygen
P

PBG
PDG
PFL
PFLD
PMG
PNEU
PNL

Pressure Breathing for g
Programmable Display Generator
Pilot Fault List
Pilot Fault List Display
Permanent Magnet Generator
Pneumatic (secondary altimeter
operating mode)
Panel
Change 1Glossary 3

T.O. GR1F16CJ1pph, PPH
PRE
PRESS.
PRI
PSA
psi, PSIPTO
PWR

Pounds per Hour
PresetPressure, Pressurization
Primary
Pneumatic Sensor Assembly
Pounds per Square Inch
Power Takeoff (shaft from engine
gearbox to ADG)
Power
Q

QTY

Quantity
R

RAD
RCR
RCVV
RDR
RDY
REL
RER
RIU
RMLG
RNDS
RNG
rpm, RPM
RS
RSE
RSVR
RSVRS
RT
RV
RWD
RWR
RZIP

Radio (e.g., RAD 1 or RAD 2)
Runway Condition Reading
Rear Compressor Variable Vanes
Radar
Ready
ReleaseRadial Error Rate
Remote Interface Unit
Right Main Landing Gear
Rounds (Gun)
Ranging
Revolutions per Minute
Return to Search
Reduced Speed Excursion
Reservoir
Reservoirs
RetardedReceive Variable
Right Wing Down
Radar Warning Receiver
Rendezvous Initial Point
S

SAI
SD
SEC
SEL
SEAWARS
SFO
SIF
SL
SMS
SNSR
SPD BRK
SPL
SQ
SQL
STA
STAPACSTBYSTDST STA

Standby Attitude Indicator
Start Descent Point
Secondary Engine Control
SelectSeawater Activated Release System
Simulated Flameout Landing
Selective Identification Feature
Sea Level
Stores Management System
SensorSpeedbrakeSound Pressure Level
SquareSquelchStationStabilization Package
StandbyStandard
Stores Station

Glossary 4Change 1

SUU
SW
SYM
SYS

Suspension Utility Unit
SwitchSymmetrical
System
T

TACAN
TAS
TBRF
TCN
TCTO
TEF's
TEMP
TER
TEU
TGM
TGT
THEO
T.O.
TOD
TR, T/R
TRV
TT
TV
TVS
TWS

Tactical Air Navigation
True Airspeed
Tunable Band Reject Filter
TACAN
Time Compliance Technical Order
Trailing Edge Flaps
Temperature
Triple Ejector Rack
Trailing Edge Up
Training Guided Missile
TargetTheoryTakeoff
Time of Day
Transmit/Receive
TravelTotal Temperature
Television
Television Sensor
Threat Warning System
U

UFC
UHF
UNK

Upfront Control
Ultra High Frequency
Unknown
V

VAC
VDC
VHF
VIP
VMC
VMS
VOL
VSV
VVI

Volts acVolts dc
Very High Frequency
Visual Initial Point
Visual Meteorological Conditions
Voice Message System
VolumeVariable Stator Vanes
Vertical Velocity Indicator
W

W/
WB
W/O
WOD
WOW
WPN
WPN RELwt, WTWith
WidebandWithout
Word of Day
Weight on Wheels
WeaponWeapon(s) Release
WeightYY

Yaw

T.O. GR1F16CJ1TERMS/SYMBOLSAn

Normal Acceleration (G's)σ

Density Ratioaο

Speed of Sound at Sea Level,
Standard DayβAngle of Sideslip

&

And

Fe

Elevator Stick Force (lb).
fRoll Rate (Deg/Sec)

Fa

Aileron Stick Force (lb).
yYaw Rate (Deg/Sec)

Fp

Rudder Pedal Force (lb)Ay

Lateral Acceleration (Ft/Sec2)

N2

Engine Compressor RPM or
Nitrogen.
q

Pitch Rate (Deg/Sec)P

PressureαAngle of Attack (Deg)T

Absolute Air TemperaturePο

Pressure at Sea Level, Standard
Day

W

Specific Weight (Lb/Ft 3)

PtTotal Pressure

Vc

Calibrated Airspeed

Ps

Static Pressure or Specific Energy
Rate

Vt

True Airspeed∆ΗDelta Change Altitude (Ft)qc

Impact Pressure (Pt - Ps).
H

Altitude Rate (FPS)ρο

Density at Sea Level, Standard Day

Glossary 5/(Glossary 6 blank)